czech aerospace - Výzkumný a zkušební letecký ústav

Komentáře

Transkript

czech aerospace - Výzkumný a zkušební letecký ústav
CZECH
AEROSPACE
Proceedings
Brno
2007
LETECK Ý
zpravodaj
In this issue:
Czech Aerospace
Research Centre
CLKV
Proceedings of the
7th Annual
Workshop held at
Brno, Czech
Republic
November 8 to 9,
2007
Centrum leteckého
a kosmického
výzkumu
CLKV
Sborník vybraných
referátů
přednesených na
7. ročníku
semináře CLKV
E
CK
Ý Ú
S
TA
T
Brno, 8. — 9.
listopadu 2007
V
U
T
O
V
LE
N o v e m b e r
2 0 0 7
ISSN 1211—877X
B
RN
No. 3 / 2007
CZECH
AEROSPACE
P r o c e e d i n g s
J OU R N A L
F O R
C Z E C H
AE RO S PAC E
R E S E A R C H
LETECK Ý
zpravodaj
VÝZKUMNÝ
Ý
Ý A ZKUŠEBNÍ
Š
Í LETECKÝ
ÝÚ
ÚSTAV, a.s.
Editorial address:
Aeronautical Research and Test Institute / VZLÚ, Plc.
Beranových 130, 199 05 Prague 9, Letňany
Czech Republic
Phone.: +420-225 115 223, Fax: +420-869 20 518
Editor-in-Chief:
Editor & Litho:
Ladislav Vymětal (e-mail: [email protected]
y
)
Stanislav Dudek (e-mail: [email protected])
Editorial Board:
Chairman:
Vice-Chairman
Members:
Publisher:
Printing:
Milan Holl, President ALV, Managing Director VZLÚ
Vlastimil Havelka, ALV
Jan Bartoň, Tomáš Bělohradský, Vladimír Daněk,
Luboš Janko, Petr Kudrna, Pavel Kučera, Oldřich Matoušek,
Zdeněk Pátek, Antonín Píštěk
Czech Aerospace Manufacturers Association / ALV, Prague
Studio Winter Ltd. Prague
Subscription and ordering information available at the editorial address. Legal liability for published
manuscripts’ originality holds the author. Manuscripts contributed are not returned automatically to
authors unless otherwise agreed. Notes and rules for the authors are published at our Internet pages
http://www
p
.vzlu.cz/.
Czech AEROSPACE Proceedings
Letecký zpravodaj
3/2007
© 2007 ALV / Association of Aviation Manufacturers, All rights reserved. No part of this publication may
be translated, reproduced, stored in a retrieval system or transmitted in any form or by any other means, electronic,
mechanical, photocopying, recording or otherwise without prior permission of the publisher.
ISSN 1211 - 877X
1
L E T E C K Ý Z P R AV O D A J
3/2007
Contents / Obsah
2
Aerodynamic Design of V48 Model Propellers
Aerodynamický návrh modelových vrtulí V48
7
Ing. Jan Dostál, Ing. Pavel Klínek
Measurement of Tension and Torque Moment of Propeller
Měření tahu a krouticího momentu vrtule
10
Ing. Ivo Jebáček
Application of Artificial Neural Networks for Searching of Faulty Components
of Turbine Engine
Použití neuronových sítí k nalezení poškozené části turbinového motoru
15
Ing. Jaromír Lamka, CSc.
Evaluation Methodology of RTD projects— chapter Verification of Evaluation
Methodology at Specific Projects
Hodnoticí metodika projektů výzkumu a vývoje — kapitola Verifikace modelů na konkrétních projektech
18
Ing. Klára Grammetbauerová, Ph.D.
Analysis of Fatigue Crack Growth under the Spectrum Loading
Výpočet šíření trhliny při zatěžování spektrem
22
Ing. Petr Augustin, Ph.D.
Measurement and Evaluation of the Interior Noise in Aircraft
Měření a hodnocení vnitřního hluku v letadlech
27
Ing. Tomáš Salava, Dr.Sc., Ing. Marcela Šloufová
Composite Recycling — Technology, Recycles, Their Parameters and Possible Applications
Recyklace kompozitů — technologie, recykláty, jejich parametry a možné aplikace
32
Ing. Miroslav Valeš, Ing. Bedřich Štekner, Karel Cihelník, Ing. Jan Grégr, Ing. Vladimír Kovačič
New FSW Equipment at VZLÚ, Plc.
Nové vybavení pro třecí svařování ve VZLÚ, a.s.
35
Ing. Petr Bělský
Buckling of Shells
Stabilitní úlohy skořepin
38
Ing. Karel Patočka, Ing. Tomáš Jamróz, Jiří Had
Numerical Study of Steady and Unsteady Flow in a Centrifugal Compressor
Porovnání stacionárního a nestacionárního řešení odstředivého kompresoru
40
Ing. Jan Tůma
Industrial Measurements of Frequency Characteristics of Small Sport Aircraft
Průmyslová měření frekvenčních charakteristik malých sportovních letounů
42
Ing. Karel Weigel, Ing. Tomáš Kostroun, Doc. Ing Svatomír Slavík, CSc
Optimalization of Stiffened Panel — Design of Testing Equipment
Optimalizace vyztuženého potahu — návrh zkušebního zařízení
46
Ing. Miroslav Pešák, Prof. Ing. Antonín Píštěk, CSc.
Design and Manufacture of Aircraft Parts by LF Technology
Konstrukce a výroba součástí letadel technologií LF
49
Ing. (Chem. E.) Lukáš Křípal
Ducted Fan Power Unit Demonstrator for Ultra Light Airplanes
Demonstrátor ventilátorového pohonu pro ultralehká letadla
52
Ing. Pavel Růžička
The Methodology of Winglet Aerodynamic Design for UL and VLA Aircraft
Metodika aerodynamického návrhu wingletu pro UL a VLA letoun
55
Ing. Robert Popela Ph.D, Ing. Pavel Zikmund, Ing. František Vaněk Ph.D, Ing. Martin Kouřil Ph.D
Exchanger Integral Property Computation and CFD Simulation of an Unconventional
Cooling System
Výpočet integrálních charakteristik nekonvenčního chladicího systému
Ing. Erik Ritschl, Jian Chen, M.Sc.
58
Shear Buckling Analysis of Composite Sandwich Panels
Ztráta stability kompozitových sendvičových panelů při smykovém namáhání
62
Ing. Martin Baumruk, Ing. Ivan Jeřábek, Ing. Karel Barák
Modelling of the Microaccelerometer MAC Translation Control - Measurement Channel
Modelování translačního kanálu řízení (měření) mikroakcelerometru MAC
65
Ing. Viktor Fedosov
Impact Analysis of the Rigid Body on the Thin-walled Aluminum Structure with
Considering of the Stochastic Material Properties
Simulace nárazu tuhé trubky na tenkostěnnou duralovou konstrukci s experimentem
Michal Mališ
2
C Z E C H A E R O S PA C E P R O C E E D I N G S
Aerodynamic Design of V48 Model
Propellers
Aerodynamický návrh modelových vrtulí V48
Ing. Jan Dostál, Ing. Pavel Klínek / VZLÚ, Plc., Prague
Aerodynamic design procedure of an advanced model propeller is described in the contribution.
contribution The propeller named V48 is
designed as an alternative propeller for the EV-55 twin-engine turboprop. The selection of the propeller diameter of 2m and the
blade number 6 has been done in the light of high propulsive efficiency and quiet aerodynamic noise. The blade airfoil family
complies with the blade tip Mach numbers which reach 0.75 under cruise design conditions. Ten blade airfoil aerodynamic
characteristics calculated for cruise conditions are described. Partial aerodynamic optimization done for the blade aspect ratio and
twist is described. V48 propeller induced velocity calculated components and acoustic pressure levels are graphed. In order to
verify design calculations by taking aerodynamic measurements in a VZLU 3m dia wind tunnel, the V48 propeller was scaled to
a model of 1 m dia, which was named ”V48“ model, and tested in the tunnel, meeting the requirements of CLKV-C3 research
project.
V příspěvku je popsána metodika aerodynamického návrhu pokročilé modelové vrtule V48, zvolené jako alternativní vrtule pro
dvoumotorový turbovrtulový letoun EV-55. Průměr 2m a počet listů 6 navrhované vrtule je vybrán z hlediska kompromisu mezi
vysokou propulzní účinností a nízkým aerodynamickým hlukem vrtule. Profilová řada pro vrtulový list je přizpůsobena Machovu
číslu na špičce listů, které při návrhovém cestovním režimu dosahuje hodnoty 0,75. Aerodynamické charakteristiky deseti profilů
listu, spočtené pro návrhový režim, jsou vykresleny pomocí kobercových grafů. Štíhlost a zkroucení vrtulového listu je optimalizováno pro návrhový režim, postup je popsán. Spočtené složky indukované rychlosti jsou vykresleny pro sedm režimů vrtule, hladiny akustických tlaků pro dva cestovní režimy. Pro plánované ověření návrhové metodiky měřením modelové vrtule v aerodynamickém tunelu VZLÚ o průměru 3m byla vrtule V48 zmenšena na průměr 1,0m a tato varianta pojmenována ”V48model“.
Keywords: propeller design, aerodynamics, induced velocity, airfoil characteristics
1. Introduction
2. V48 Diameter and Number of Blades Selection
Aerodynamic and acoustic software development and application
during the designing of two model propellers are the objectives of the
CLKV-C3 research project ”Advanced Aerodynamic Methods on
Design of Propellers“. The first of the two propellers is the V44
model propeller described in [1] and [2]. A proposal of aerodynamic
design methodology and verification of new propeller airfoils family
were aimed at the first model. The proposed methodology is used, still
upgraded and improved, in the course of designing the second, advanced model propeller named V48.
The V48 propeller is designed as an alternative propeller for the
EV-55 twin-engine turboprop aircraft. Supposed engines are
AI-450TP, made in the Ukrainian Ivchenko-Progress company. The
original EV-55 propellers are Avia AV-844 or Hartzell HC-E4N-3I,
both four-bladed, made of aluminum alloys. The V48 proposed
material is composite. The technical documentation of any of the two
EV-55 propellers is not available, thus the V48 design must be begun
afresh as contrasted to the first model propeller. Only EV-55 propeller
spinner engineering drawing succeeded to obtain, thus the blade hub
geometry complies with the spinner.
After finishing the aerodynamic design the additional development
of V48 propeller was stopped. The blade and hub construction, stiffness and dynamic calculations and experiments make the development project possible to continue if it is necessary.
The geometric dimensions of V48 propeller are scaled and the
resulting scaled propeller has been named ”V48model“. The scaled
model of V48 will be manufactured instead of the original V48 one.
The purpose of the scaling is to fit the dimensions of V48model to the
measuring area of the VZLU wind tunnel and to the torque dynamometer used today. The aerodynamics tests of the second model propeller are scheduled at CLKV-C3 research project to verify the aerodynamic design calculations.
So the second CLKV-C3 model propeller is represented by the two
model propellers: V48 and V48model.
Only the elementary data were known on mentioned Avia and Hartzell propellers before the V48 design was started (see Tab. 1). Therefore the old well-tried V510 propeller [3] served for the first design
steps. The V510 blade was constringed into three different diameters
and its depth c was alternated per multiplying with scaling coefficient
such that 15 alternatives of V510 blade were made. The each alternative number of blades was altered between 4, 6 and 8 values to increase the total propeller alternative count into 45. Even values are chosen for blades static balance reasons. Then the propulsive efficiency
calculation under course conditions was carried out for each of the
V510 propeller alternatives with the using of the Afrsx2 propeller
aerodynamics solver [4]. The results of the preliminary calculations
are pictured on Fig. 1. The scaling coefficient dependences are distorted by simplified computational model using the only airfoils aerodynamic characteristics file. A more realistic curve looks like Fig. 9 one.
However propeller diameter dependences and number of blade ones
are believed to be reliable enough. The solid lines correspond to
a maximal diameter under consideration; dashed ones correspond to
the smaller diameter and to the minimal one. Identical colors correspond to a same number of blades.
The Six V510 alternatives with greatest propulsive efficiency
advanced to the second round of calculations, this acoustic one. Maximal propeller aerodynamic noise levels on crossing at 3 km altitude,
407km/h velocity (cruise condition) are depicted on Fig. 2. Those
were calculated with the using of the propeller acoustics computer
program Aav_m3a7 [5]. Points of dashed lines correspond to acoustic
filter of A type. It stands to reason, from Figs. 1 and 2, the greater
aerodynamic losses and the weight of 8 blade propeller can not be
compensate by acoustic advantage, whereas 4 blades noise level isn't
sufficiently compensated by better propulsive efficiency. So the choice of 6 blades is definite.
The maximum diameter of 2,09m was excluded from the third
round of calculations because its aerodynamic advantages are bound
3
L E T E C K Ý Z P R AV O D A J
with 4 blades excluded in the previous item. From remaining diameters 2m alternate advanced unambiguously. Alternate 1,9m diameter
has (hub diameter)/(tip diameter) ratio at given hub diameter much
greater then the standard value of 0,2 used in the simplified aerodynamic computational model, so that its real propulsive efficiency will
be less then the calculated one. Therefore 1.958m diameter alternates
with 1.2times greater hub/tip ratio then the standard one, was the
second and last member choosing for the third round of calculations.
These calculations begun with propulsive efficiency calculation of
1.958m alternate. Then the optimal blade twist (see section 5) was
found for the both third ring alternates and efficiency and the noise
level calculation followed on. The results on Figs. 1, 2 typified by greater symbols reflect the slight efficiency increase and the noise level
decrease caused by the blade twist optimization. The definite V48 diameter value of 2m has been chosen.
Tab. 1 — Main propeller technical data comparison
3/2007
3. V48 Propeller Geometry Proposal
Primary V48 composite blade dimensions were set by the following
way:
The blade airfoil relative thickness distribution responds to V09
propeller composite blade. The blade depth (airfoil chord length) is
nearly constant along a better part of the radius; it decreases at the tip
region. Airfoil decision points lie at the blade axis for the most part of
the blade length. The hub area was suited to EV-55 — AI-450TP
spinner; the tip was turned downstream to decrease the aerodynamic
noise of the blade tip in agreement with [6].
A new airfoil family V4modvm2.trh was designed for the V48
blade airfoils. It is based on older Benda's supercritical airfoil family
V4mod. The airfoil chamber — relative thickness dependence was
changed by decreasing the tip chamber to 2% value, in addition to
smoothing family's definition constant. A new algorithm was put up
for the transitional surface from the ending family airfoil to the hub
circle. It ensures a continuous surface passing in the place of the
ending boundary (usually 20%) airfoil. This requirement was not executed at previous propeller designs.
Fig. 1. V510 Alternates Efficiency
0,88
D=1,9m; 4blades
D=1,9m; 6blades
D=1,9m; 8blades
D=2m; 4blades
0,86
D=2m; 6blades
D=2m; 8blades
D=2,09m; 4blades
0,85
D=2,09m; 6blades
Fig. 3 — A blueprint of V48 propeller blade
D=2,09m; 8blades
0,84
D=1,96m; 6bl. optim.
D=2m; 6bl. optim.
0,83
0,8
0,85
0,9
0,95
1
Scaling Coefficient c/c510
1,05
Fig. 4 — A rendering
of the V48 propeller
Fig. 2. V510 Alternates Noise Level
L Total Level
LA Filtered Level
59
58
57
56
55
54
L - 4blades
53
L - 6blades
52
Noise Levels [dB]
Propulsive Efficiency
0,87
51
L - 8blades
50
LA - 4blades
49
48
LA - 6blades
47
LA - 8blades
46
45
L - 6blades
optim.
LA - 6blades
optim.
44
43
42
41
40
39
1,8
1,9
2
Propeller Diameter [m]
2,1
4. V48 Blade Airfoil Aerodynamic Characteristics
The airfoil aerodynamic characteristics at the ten standard blade sections were computed for the cruise flight condition by the same technique and with the same solver that was used in case of the first
CLKV model propeller design [1]. The Swshl7 solver routinely calculates lift coefficients and drag (including wave drag) ones. Maximum incoming flow Mach number 0.74 falls on the tenth blade airfoil, near the blade tip. The first and only, 35.7% thick airfoil characteristics had to be interpolated between the second airfoil
characteristics and the hub circle ones. Wind tunnel experiments [7]
of the two ultimate V4mod family airfoils (8% and 10%) were used
for computed results comparison, eventually completion, range extension and correction.
Carpet diagrams (Figs. 5 to 8) are used, as well as in [1],
for CL(α), CL(r), CD(α), CD(r) dependences depiction. Mutual shifts of
individual characteristics scale factors are fixed by the delta values
brought out below the title of each diagram. CL (α) characteristics (Fig.
5) are corrected for rotation effects: Coriolis forces, which act on blade
boundary layer, protect its separation and thus delay airfoil stall condi-
4
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 5. CL(Alfa) Characteristics
1-37,7%
Delta CL = 0.4
Fig. 6. CL(r) Characteristics
Delta CL = 0.2
6
2-23,8%
-8°
-7°
4-15,9%
-6°
-5°
5
6-10,9%
-4°
7-9,0%
-3°
8-7,6%
-2°
9-6,7%
4
10-6,2%
3
-10°
3-19,0%
5-13,2%
4
-13°
-1°
0°
Lift Coefficient CL
Lift Coefficient CL
+1°
+2°
3
2
+3°
+4°
+5°
2
+6°
+7°
1
+8°
+9°
1
+10°
+12°
0
+15°
0
+19°
-1
-1 -1 -9 -7 -5 -3 -1 1
3 1
3
5
7
9 11 13 15 17 19
-1
160 240 320 400 480 560 640 720 800 880 960
Angle of Attack alfa [°]
Blade Radius r [mm]
Fig. 7. CD(Alfa) Characteristics
Delta CD = 0.02
Fig. 8. CD(r) Characteristics
Delta CD = 0.01
1-37,7%
2-23,8%
-3°
-2°
-1°
0°
+1°
+2°
+3°
+4°
+5°
+6°
+7°
+8°
+9°
+10°
+12°
+15°
+19°
-4°
-5°
-6°
-7°
-8°
-10°
-13°
0,5
3-19,0%
0,5
4-15,9%
5-13,2%
6-10,9%
7-9,0%
0,4
0,4
8-7,6%
9-6,7%
Drag Coefficient CD
Drag Coefficient CD
10-6,2%
0,3
0,3
0,2
0,2
0,1
0,1
0
0
-13 -11 -9 -7 -5 -3 -1
1
3
5
7
9 11 13 15 17 19
Angle of Attack alfa [°]
160 240 320 400 480 560 640 720 800 880 960
Blade radius r [mm]
5
L E T E C K Ý Z P R AV O D A J
Fig. 10. Axial Components of Propeller V48
Induced Velocity
Fig. 9. Blade Aspect Ratio Optimization
0,867
Single
Aerodynamic
Characteristics
0,866
Variable
Aerodynamic
Characteristics
0,865
0,863
Velocity Va[m/s]
Propulsive efficiency Eta
0,864
0,862
0,861
0,86
0,859
0,858
0,857
0,856
0,855
0,9
0,95
1
1,05
3/2007
1,1
c/c0
42
40
38
36
34
32
30
28
26
24
22
20
18
16
14
12
10
8
6
4
2
0
-2
-4
-6
-8
-10
-12
-14
-16
-18
-20
-22
Take off Vinf=0km/h
Climb Vinf=155km/h
Cruise Vinf=407km/h
Maximal Vinf=408km/h
QuietCruise Vinf=402km/h
OfAxisFall Vinf=145km/h
Reversal Vinf=150km/h
0 0,1 0,2 0,3 0,4 0,5 0,6 0,7 0,8 0,9 1
Blade Radius r [m]
tions [8]. CL (r) characteristics (Fig. 6) are somewhat more gentle then
V44 ones [1], which corresponds with the similar pattern of camber (r)
dependences. CD (α) characteristics (Fig. 7) are much more symmetrical towards α axis then V44 ones because of reversal propeller conditions are expected in this case. CD(r) characteristics (Fig. 8) in α range
from -3° to +19° are illustrated by thick lines; completive ones in
α range from -4° to -13° are illustrated by thinner lines.
All ten blade airfoil aerodynamic characteristics are coded into
text-file, which is an input data file of the Afrsx2 propeller aerodynamic solver.
one, for each of nine computed polar lines. The circle — second airfoil interpolation was then used for remaining first cut airfoil polar
line. The results of V48 blade aspect ratio optimization are etched in
Fig. 9. The dashed line on the Fig. responds to single, primary airfoil
characteristics input file of Afrsx2 solver used, the decisive solid line
responds to the proper airfoil characteristics input files obtained by
the process listed above, for each of the seven optimizing steps, given
by a c/c0 ratio. The blade aspect optimization result is insignificant
propeller efficiency increase. A suitable choice of the primary blade
aspect ratio was confirmed by the optimization.
5. V48 Partial Aerodynamic Optimization
6. V48 Geometric Blade Form Smoothing
The first part of the V48 propeller aerodynamic and acoustic optimization is described in section 1 within the context of the propeller diameter and number of blades search. The methodology used for
a blade twist and aspect ratio optimization is inscribed in [1], [2]. The
same proceedings were used for the V48 propeller. The single-point
optimization covered again only one, cruise flight condition. The specific cost function is the propeller propulsive efficiency again.
When the blade aspect ratio is optimizing, the blade diameter stays
constant, the blade depth or airfoil chords are changed in such a way
that ratio c/c0 is constant along the whole blade; where c are new
chord lengths and c0 are chord lengths before the optimization begun.
Airfoil absolute thicknesses stay constant over the whole optimization process, the relative thickness increases as the chord decreases and
vice versa. Reynolds numbers of all ten aerodynamic cuts airfoils
change together with theirs chord change, so new airfoil aerodynamic
characteristics CL (α), CD (α), which correspond to actual relative
thicknesses and Reynolds numbers, are needed. These were obtained
per calculation for only three angles of attack α near by the cruise
condition angles and per each characteristics curve shift about difference between actual average appropriate coefficient value and primer
A smooth, fluent blade surface is ensured partly by airfoil family constant smoothness and fluency, partly by blade dimensions ones. Blade
dimensions Afrsx2 input text-file is tabular dependence of five definition dimensions on the blade radius. Those are: airfoil relative thickness,
chord length, leading edge x- and y-dimension and airfoil chord stagger
angle. Smoothing of these parameters using the technique described in
[1], [2] presents last V48 propeller aerodynamic development work.
A few operating condition aerodynamics including strength and stiffness of homogenous material (duralumin) blade computations were
done before the smoothing began. Radii of maximum strength were
kept in mind and when smoothing calculations run out, care-take was
done for thickness and chord discrepancies from primary dimensions at
those radii — only positive values are acceptable.
7. V48 Induced Velocity Components along the
Blade Length
V48 blade induced velocity components calculations were done for
the 7 operating conditions (Fig. 10 to 12). In comparison with V44
propeller [1] two other conditions swell. Those are quiet cruise and
reversal. The third new condition, off axis fall, was included into V44
calculations, but it wasn't included into [1] article. The cruise, quiet
6
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 11. Circumferential Components of
Propeller V48 Induced Velocity
20
18
16
14
12
10
10
8
6
4
2
0
-2
Velocity Vr [m/s]
Velocity Vt [m/s]
12
Take off Vinf=0km/h
Climb Vinf=155km/h
Cruise Vinf=407km/h
Maximal Vinf=408km/h
QuietCruise Vinf=402km/h
OfAxisFall Vinf=145km/h
Reversal Vinf=150km/h
8
6
Fig. 12. Radial Components of Propeller V48
Induced Velocity
-4
-6
-8
-10
-12
4
Take off Vinf=0km/h
Climb Vinf=155km/h
Cruise Vinf=407km/h
Maximal Vinf=408km/h
QuietCruise Vinf=402km/h
OfAxisFall Vinf=145hk/h
Reversal Vinf=150km/h
-14
-16
2
-18
0
-20
-22
-2
-24
-4
-26
0 0,1 0,2 0,3 0,4 0,5 0,6 0,7 0,8 0,9 1
0
0,1 0,2 0,3 0,4 0,5 0,6 0,7 0,8 0,9
Blade Radius r [m]
Blade Radius r [m]
cruise and maximal velocity conditions do not vary too much, so their
graphs almost float. Each of operating conditions is brought on graph's legend together with a relevant air speed.
Maximum induced velocity component values can be found for axial
components at the take off condition (Fig. 11). These reach more than
40 m/s, the great value corresponds to the relative great number of blades. Magnitudes of induced velocity components decrease with the air
speed; those magnitudes increase with propeller revs and with engine
power, simultaneously. Positive axial direction component aims against
the aircraft air speed in all conditions except reversal one.
Circumferential induced velocity components (Fig. 11) are lossmaking. These have less magnitudes then axial components, fortunately. Positive circumferential component direction aims with a propeller
rotation in all conditions except reversal one, again.
Radial induced velocity components (Fig. 12) change their direction
inside the propeller disc. Their positive direction (oriented in a propeller
radius) at a hub and negative one at a tip means contraction of the propeller stream (in all conditions in view except reversal one again) and on
the contrary, negative direction at a hub and positive one at a tip means
expansion of propeller stream in reversal or windmill condition.
Fig. 13. V48 acoustic pressure levels when
crossing at 10000ft
60
55
Acoustic pressure level L, LA [dB]
50
45
40
35
30
25
20
8. V48 Propeller Noise
L Cruise
15
L Quiet cruise
10
LA Cruise
LA Quiet cruise
5
0
30
50
70
90
1
110
Emission angle [deg]
130
150
The noise of propellers has the same importance like their efficiency
nowadays. Hence it was calculated in early stage of the design when
the propeller diameter dimension and number of blades were looked
for. (See section 2, Fig. 2).
The checking calculations of acoustic pressure levels were done for
the two cruising conditions in the final stage of the design. The results
of the calculations with the using the propeller acoustics computer program Aav_m3a7 [5] again are displayed on the Fig. 13. Both conditions
have the same altitude, 10 000ft (3048m). The first of the conditions
(”Cruise“) was the condition of the partial aerodynamic optimization
(see section 5). It is determined with the 407km/h TAS flight velocity,
7
375kW engine power, and 2150 rev/min propeller speed. The second
one (”Quiet cruise“) is determined with 402km/h TAS flight velocity,
364kW engine power, and 1950rev/min propeller speed.
The Acoustic pressure level L is the total sum of single frequency
parts along the whole frequency range. Acoustic pressure level LA is the
total sum of single frequency parts corrected by the acoustic filter of the
A type [9] which puts resultant values near human ear perceptions.
Relative differences between separate condition values on the Fig.
13 are much greater than induced velocity components ones on the
Figs. 10 to 12.
L E T E C K Ý Z P R AV O D A J
3/2007
This article deals with the two second model propellers. The first of
them, six bladed composite V48, could replace the four bladed metal
propellers of the Evektor EV-55 aircraft. The V48 propeller(s) would
be quieter than the Avia and Hartzell ones with comparable propulsive
efficiency. But the model propeller designed within the scope of the
research project is not intended for commercial purposes.
References:
[1]
Dostál, J., Klínek, P.: Aerodynamic Design of V44 Model Propeller
and Aeroacoustic Characteristics Calculation; Czech Aerospace
Proceedings, Vol. 2005, No. 3, November 2005, pp. 34-39
[2]
Dostál, J., Hraška, M., Klínek, P.: Aerodynamický návrh základní
modelové vrtule V44 a konstrukce listu; VZLÚ Report R-3909
(2006), p. 92
[3]
Benda, L.: Podklady pro projekt vrtule V510; VZLÚ Research
Report V-1497/83, p. 150
[4]
Dostál, J.: Výpočtové programy aerodynamiky vrtulí podle vírové
metody radiální nosné úsečky; VZLÚ Report R-3656 (2004), p. 34
[5]
Klínek, P.: Manuál k programu Aeroakustická Analýza vrtule;
VZLÚ Report R-3891 (1976), p. 25
[6]
Wagner, S., Bareiß, R., Guidati, G.: Wind Turbine Noise; SpringerVerlag Berlin Heidelberg (1996), p. 204
[7]
Benda, L., Dostál, J.: Aerodynamické profily pro kompozitové rotující nosné plochy; Závěrečná zpráva; VZLÚ Research Report
V-1685/97, p. 124.
[8]
Gur, O., Rosen, A.: Propeller Performance at Low Advance Ratio;
Journal of Aircraft, Vol. 42, No. 2, March-April 2005, pp. 435-441
10. Conclusion
[9]
The CLKV-C3 model propellers are the flash points of the research project. Those are a touchstone of the project simultaneously. Any errors or
mistakes whatsoever made by the authors will come to light during scheduled tests, whether they will be done on an aircraft - or in a wind tunnel.
Madejewski, B.: Aeroakustika, základy teorie a aplikace na konstrukci letadel; Rektorát Vysokého učení technického v Brně,
1986, p. 76
[10]
Hartman, E., P., Biermann, D.: The Aerodynamic Characteristics of
Full-Scale Propellers Having 2, 3, and 4 Blades of Clark Y and R.
A. F. 6 Airfoil Sections; NACA Report No. 640 (1938), pp. 547-569
9. The V48model Propeller
All V48 propeller aerodynamic design data are stored and further constructional work can continue where it will be appropriate.
The aim of the CLKV-C3 research project is experimental verification of aerodynamic calculation among others. Wind tunnel experiments
are scheduled in the VZLÚ 3m wind tunnel. It was necessary to scale
the designed propeller for this purpose. The 0.75 scaling factor was
determined in agreement with Hartman and Biermann [10], their old
article is frequently quoted at contemporary technical papers. But torque dynamometer difficulties require other scaling factor reducing to
the 0.5 value. The scaled V48 propeller is named ”V48model“. Hence
the second CLKV model propeller is represented by the two designed
propellers: V48 and V48model. The last of them is intended for manufacture.
Measurement of Tension and Torque
Moment of Propeller
Měření tahu a krouticího momentu vrtule
Ing. Ivo Jebáček, PhD. / Institute of Aerospace Engineering, Brno University
of Technology
The article deals with measurement of tension of propeller during flight of the KP-2U
KP 2U Owl aircraft using strain gauges.
gauges First of all,
all
methods of this tension measurement at an engine mount sample were verified under laboratory conditions, and then the strain
gauges were installed on the engine mount of aircraft. After calibration, the measuring flights were done and used for determination
of tension moments of the propeller for different time regimes. These results are used to set polars of KP-2U Owl aircraft.
Příspěvek se zabývá měřením tahu vrtule za letu na letounu KP-2U Sova s využitím tenzometrů. Nejprve byly laboratorně ověřeny
možnosti tohoto způsobu měření tahu na vzorku motorového lože a poté byly tenzometry instalovány na motorové lože letounu. Po
kalibraci byly provedeny měřicí lety z nichž byly stanoveny tahy vrtule pro různé režimy letu. Tyto výsledky slouží ke stanovení
poláry letounu KP-2U Sova.
1. Introduction
The goal of these measurements is to set the force and torque
moments of the propeller with the aid of deformation measurement at
an engine mount. The engine mount from aircraft KP-2U Sova (Owl)
was used for the first test under laboratory conditions. Figure 1 shows
the diagram of the engine mount.
2. Laboratory testing
Strain gauges were installed on each pipe as half bridges for strain
gauges measurement. Influence of temperature was excluded. A possible bend of individual pipes would not be excluded. In this case we
do not want to set exact tension of pipes in engine mount. We need to
set the most linear thrust transfer and torque moment on deformation
and this is why the possible bend of pipes is desirable because it will
help to increase sensitivity of the bridge on given load.
Standard foil strain gauges were used for these purposes. The
solid-state strain gauges could be installed, in case these strain gauges
had small sensitivity. Fortunately, first tests proved sufficient sensitivity of the foil strain gauges. Total number of installed strain gauges
is 8.
Force was inserted on the propeller during calibration. This force
together with responses of strain gauges bridges were recorded by
ESAM sensor. Torque moment was deduced similarly.
8
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 1 Diagram of engine test bed mount of KP-2U aircraft and
space for strain gauges installation
Considering the installation of strain gauges on 4 main pipes of engine mount (Fig. 1), calibration equation is:
m1
T
MK
=
b11
b12
b13
b14
b21 b22
b23
b24
*
m2
m3
m4
where b is searching calibrating coefficients;
m1 to m4 are responses to strain gauges;
T and M are applied calibrating loads.
Values b coefficients for propeller thrust are given by equation:
T
b ij = m nj * m nj
−1
T
* m nj * T
Fig. 3 Engine mount for laboratory testing
This prepared engine mount was calibrated in the laboratory, strain
gauges were installed and according to Fig. 1 calibrating equation for
engine thrust was set:
T = (0,309 ± 0,078)m1 + (0,328 ± 0,004 )m2
+ (− 2,098 ± 0,2630 )m3 + (− 1,68 ± 0,175 )m4
and for torque moment:
M K = (0,235 ± 0,048)m1 + (− 0,309 ± 0,015 )m 2
3. In-flight measurements
On the basis of laboratory experiences it was possible to conduct
experiments at the engine mount on specific KP-2U Sova aircraft.
That was why 4 strain gauges were installed at the engine mount
(see Fig. 4).
It will be possible to set a probable error of calibrating coefficients
and then the final engine thrust force will be obtained from equation:
T = (b11 ± P.E.(b11 ) )m1 + (b12 ± P.E.(b12 ) )m 2
+ (b13 ± P.E.(b13 ) )m3 + (b14 ± P.E.(b14 ) )m4
where P.E.(bijj) is a probable error of calibrating coefficients.
The probable error of estimate of engine thrust or torque moment
values were obtained from equation:
P.E.(T ) = 0,6745.
Where
εv
n
q
∑ε
2
V
n − ( q + 1)
sum of squares of the residuals
number of loads applied
number of coefficients in calibration equation
Fig. 4 Strain gauges installed
Fig. 2 Engine mount of the KP-2U Owl aircraft
Strain gauges were independently connected into data logger.
The calibration was done while aircraft was attached with a rope
with a sensor. After the engine was started up, it was smoothly
step on the gas and force was measured on the sensor together
with strain gauges responses. Aircraft wheels were fixed on a plywood pad. Maximal force of static tension with the smallest angel
of stall and 5000 revolution per minute was about 795N. Strain
gauges responses from the right and the left part of engine mount
were count up for these calculations, and then we obtained
9
L E T E C K Ý Z P R AV O D A J
3/2007
Fig. 5 Dependence of strain gauges
responses on tension force of propeller
1000
Response [um/m]
500
0
0
200
400
600
800
1000
um/m
1200
um/m
-500
-1000
-1500
Force [N]
Fig. 5 Dependence of tension force on
speed — stable flights
1800,0
1600,0
1400,0
Force [N]
1200,0
1000,0
800,0
600,0
Fig. 6 (below) Aircraft polar set from
measured data
400,0
200,0
1,0000
0,0
80,0
100,0
120,0
140,0
160,0
180,0
Vias [km/h]
0,8000
0,7000
0,6000
Cl [-]
one calibration constant including probable mistake
b = 0,6034±0,0042. Fig. 5 shows the graph of dependence of
propeller tension and strain gauges response.
After calibration, the measuring flights were done. Fig. 6
shows measured tension force in dependence on speed. Measurement was always done after stabilizing of required speed, and
then arithmetic mean was set from measured data. Fig. 7 shows
polar that was calculated from measured data. However, aerodynamic correction of aircraft speed system was not done and the
aircraft mass was also assessed. Therefore the initiated results can
be considered only as orientation ones and we can use these data
for further experiments.
200,0
0,9000
0,5000
0,4000
0,3000
References:
[1]
[2]
[3]
Jebáček I.: Measurement of Horizontal Tail Load by Strain Gauges; Czech Aerospace Proceedings / Letecký zpravodaj, Prague, 2004
Jebáček I.: Calibration of strain-gage installation in stabilizer structures for the measurement of flight loads;
Grant Project FP 390038, TU Brno, 1999
Skopinski T.H., Aiken William S.: Calibration of straingage installation in aircraft structures for the measurement of flight loads; NACA Report 1178
0,2000
0,1000
0,0000
0,0800
0,0850
0,0900
0,0950
0,1000
Cd [-]
0,1050
0,1100
0,1150
0,1200
10
C Z E C H A E R O S PA C E P R O C E E D I N G S
Application of Artificial Neural Networks
for Searching of Faulty Components
of Turbine Engine
Použití neuronových sítí k nalezení poškozené části
turbinového motoru
Ing. Jaromír Lamka, CSc. / VZLÚ, Plc., Prague
Ai
Aircraft
ft tturbine
bi engines
i
b
become d
degraded
d dd
during
i operation
ti and
d their
th i associated
i t d maintenance
i t
costs
t can become
b
extremely high for owners. Hence, successful maintenance techniques are those which are able to reduce maintenance
costs and down-time. In recent decades predictive maintenance techniques have been used because of their benefits
in reducing down-time compared to traditional techniques like breakdown maintenance. As a result different
predictive maintenance and diagnostics techniques have been developed during the last fifteen years. This study, in
particular, will focus on performance diagnostic techniques based on Neural Networks. The network features and
training algorithms will be discussed to develop an appropriate model for gas turbine diagnostics. In addition, it will
be shown how training data can affect training performance.
Parametry leteckých turbinových motorů v průběhu provozu degradují. S tím spojené náklady na údržbu pro
provozovatele extrémně rostou. Mezi úspěšné metody údržby patří ty, které redukují jak náklady na údržbu, tak
dobu ”uzemnění“ stroje. V posledních desetiletích začaly být používány predikční metody údržby, neboť výrazně
snižují prostojové časy ve srovnání s tradičními metodami jako např. údržba na pevné časy doplněná o další
poznatky. Tento příspěvek je zaměřen na diagnostickou metodu, která se opírá o neuronové sítě. Jsou zde
diskutovány vlastnosti sítí a algoritmy učení sítí za účelem vývoje vhodného modelu pro diagnostiku turbinového
motoru. Bude ukázáno, jak množina učicích dat ovlivní efektivitu učení.
Keywords: Gas Turbine Engine, Sensor, Diagnostics, Artificial Neural Network.
Introduction
Gas turbines are used under a wide range of operational temperature,
speed, power and environmental conditions for long periods of time,
which means, the components can become degraded due to fouling,
corrosion, erosion, thermal fatigue and foreign object damage. Performance degradation results in a rise in Turbine Entry Temperature
(TET) and Specific Fuel Consumption (SFC) — to keep shaft speed
constant, reduced power and a change in compressor surge margin. If
no maintenance action is carried out, degradation results in unschedu-
Nomenclature
ANN . . . . . . .artificial neural network
BSI . . . . . . . .British Standards Institution
ICM . . . . . . . .influence coefficient matrix
EHM . . . . . . .engine health monitoring
FCM . . . . . . .fault coefficient matrix
GPA . . . . . . . .gas path analysis
MSE . . . . . . .mean square error
nG . . . . . . . . .gas generator rotor speed
nP . . . . . . . . . .propeller shaft speed
p0 . . . . . . . . . .barometric pressure
P . . . . . . . . . .output power
p2C . . . . . . . . .total outlet pressure of compressor
SFC . . . . . . . .specific fuel consumption
TET . . . . . . . .turbine entry temperature
Tk . . . . . . . . . .torque on propeller shaft
T0 . . . . . . . . . .ambient temperature
T2C . . . . . . . . .total outlet temperature of compressor
T4C . . . . . . . . .total inter turbine temperature
Wfe . . . . . . . . .fuel flow
led shut downs of the engine which can be very costly. Therefore
techniques which detect the existing component failures should be
employed for the efficient maintenance management of modem turbine engines. During the last sixty years, many changes in maintenance
techniques and management have occurred and lots of efficient maintenance techniques have been developed due to competition between
aircraft turbine engine manufacturer to minimize the engine running
costs.
The British Standards Institution (BSI) [l] has defined maintenance
as the combination of all technical and associated administrative actions intended to retain an item or system in, or restore it to, a state in
which it can perform its required function. No maintenance technique
can completely eliminate failures; however maintenance can be planned which results in the lowest possible downtime as a consequence
of failure. It means maintenance techniques are selected to produce
high plant availability, where availability has been defined by Kumar
[2] as the probability that a system or equipment, when used under
specified conditions in an ideal environment (i.e. readily available
tools, spares, personnel, etc.) will operate satisfactory at any point in
time as required.
Based on Dur new understanding of availability and failure modes,
new maintenance concepts are created. Figure 1 shows how new
maintenance tools have lowered down-time since the 1930s.
Regarding Fig. 1, diagnostics includes analytical techniques based
on gas turbine performance analysis and actual engine parameters
read from instruments. The first developed procedures like Fault
Trees and Fault Matrices are still used widely to provide qualitative
information; however they are unable to detect two or more failures
occurring simultaneously within the engine. The further developed
technique, Gas Path Analysis (GPA) can produce quantitative results
11
L E T E C K Ý Z P R AV O D A J
3/2007
Figure 2 — Walter M-601 Turboprop Engine Layout
Figure 1 — Development of new maintenance techniquees
instead of qualitative results but the GPA is a linear diagnostic technique, which is therefore limited to small degradation changes. These
limitations of the above techniques have led developers to using more
sophisticated techniques like non-linear GPA, Artificial Neural Networks (ANN), genetic algorithms and expert systems.
GPA and its derivatives are powerful diagnostic approaches, although they require certain conditions (Doel [1993]) to work well, and
they suffer from disadvantages such as the ”smearing“ effect. It has
been evident for a long time that if degraded gas turbine components
can be located with a priori information, GPA approaches can be very
successful in quantifying the degradation. This idea has been discussed by many researchers, e.g. Mathioudakis [2003].
The source of the a priori information may be different; one of
which is the fault isolation technique. Different gas turbine fault isolation techniques, such as artificial neural networks (ANN) (Ogaji and
Singh [2003]) and pattern matching method (Lee and Singh [1996])
have been developed. In this paper, the fault pattern matching method
introduced by Lee and Singh [1996] has been developed further, used
to isolate degraded gas turbine components and initially quantify the
degradation. Once the degraded components are located, non-linear
GPA is used to refine the degradation assessment. Non-linear GPA
can deal with higher levels of degradations.
Engine performance modelling
The application of artificial neural networks was aimed at diagnostics
of Walter M-601E turboprop engine (Fig. 2). The M-601E engine
falls into a category of small engines. The engine was designed in
VZLÚ and its manufacturer is the Walter company. Initially the engine was determined for Czech LET L-410 commuter. At present there
are agricultural, high altitude, and aerobatic versions. Walter engines
of the M-601 family are two-shaft turboprop aircraft engines with
a free power turbine and reverse flow configuration.
Engine ground performance under the ISA conditions:
- max. take-off powe r . . . . . . . . . . . . . 751 shaft horse power,
- max. continuous . . . . . . . . . . . . . . . . 657 shaft horse power,
- gas generator speed of rotation . . . . . . 36 660 rpm (100%),
- propeller shaft . . . . . . . . . . . . . . . . . . . . 2080 rpm (100%),
- inter turbine temperature . . . . . . . . . . . . . . . . . . . . . 735°C,
- air mass flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.6 kg/s,
- specific fuel consumption . . . . . . . . . . . . . . . 0.395 kg/kWh.
The engine was thermodynamically modelled by means of the
M-601 simulation program. The program is able to respect both installation of the engine into the nacelle and engine movement during the
flight, too.
ce the impact of measurement noise and the mathematical expression
for the averaging is shown in Equation (1).
z=
1 m
∑ zi
m i =1
(1)
A linear and non-linear GPA approach developed at Cranfield (Escher
[1995]) is an effective gas turbine diagnostic method to assess gas turbine faults. The approach is based on the assumption that a linearized
gas turbine performance model, Equation (2), is an accurate description of engine behaviour at certain operating conditions.
Δz = H ∗ Δx
(2)
The degradation of components expressed by the deviation of component parameters (flow capacity and efficiency) can be predicted by
inverting of the influence coefficient matrix (ICM) H to a fault coefficient matrix (FCM) H-1 leading to Equation (3).
Δ x = H −1 ∗ Δ z
(3)
The nonlinearity of the engine behaviour is taken into account by
using an iterative process, Escher [1995], where linear GPA is applied iteratively until a converged solution is obtained, Fig. 3.
This diagnostic method is simple and fast in prediction although it
has certain requirements that are difficult to meet such as:
● determining the ICM to accurately describe gas turbine performance,
● fault and noise free sensors,
● uncorrelated measurements that are sensitive to engine degradation.
Another disadvantage of GPA is the uncertainty of the prediction —
conflicting answers may be provided if degraded components are
unknown. Therefore, further interpretation of the prediction from
GPA may be very important in order to avoid being misled by GPA.
This job can only be done by experienced gas turbine diagnostics
engineers. To show the severity of the uncertainty, a simulated degradation and diagnostic process with GPA using the model engine is presented as follows.
Gas path analysis
In order to improve the effectiveness of gas turbine diagnostics, measurement noise reduction techniques must be applied to reduce the
impact of measurement noise before GPA is used. Many measurement noise reduction techniques, such as those described by Lu et al.
[2000] and Ganguli [2001], have been developed in the past. A discussion of these techniques is beyond the scope of this paper. Therefore, a simple data averaging method is used in this research to redu-
Figure 3 — Illustration of non-linear
non linear GPA technique where Δ⎯x is independent parameter(s) and Δ⎯z is dependent parameter(s)
C Z E C H A E R O S PA C E P R O C E E D I N G S
12
Optimum instrumentation selection
Problem is the Walter M-601 turboprop engine has no suitable number, types and location of sensors (Tab. 2). Only efficiency of compressor can be calculated in accordance with parameters mentioned in
Table 2. The fact that the propeller shaft speed and the torque on propeller shaft have mechanical bond is necessary take into account.
Strict distribution of faults listed in Table 1 can not be used in this
case. A combination of parameters for classification parts of engine
can be applied for:
compressor — . . . . . . . . . . . . . . . . . . . . . . . . . . p2C, T2C, nG,
gas generator turbine — . . . . . . . . . . . . . . . . . . . T4C, nG, Wfe,
power turbine — . . . . . . . . . . . . . . . . . . . . . . . . . Tk, nP, Wfe.
Monitoring dependent (i.e. measured) parameters for Engine Health
Monitoring (EHM) requires instrumentation that in some cases can be
complex and highly costly. A lower number of monitored dependent
parameters reduces technique complexity and costs but can reduce
the accuracy of the detected engine faults. Hence, a successful EHM
technique requires optimum selection of dependent parameters. The
word optimum in this paper reflects the suitable number, types and
location of sensors.
When a failure occurs it changes independent parameters like efficiency and flow capacity. However, independent parameters are not
measurable, but changes in independent parameters produce changes
in measurable dependent parameters. Therefore, identification of
changes in dependent parameter will allow fault isolation and detection if the aero-thermodynamic relationships are known. Non-linear
GPA is an improved technique of linear GPA. To make non-linearity,
each performance change is added to the last performance characteristic and the process is iterated to convergence as illustrated in Fig. 3.
Compressor fouling and turbine erosion have a high probability of
occurrence in a turbine engine, therefore the following independent
parameters are usually considered as part of the instrumentation
study:
a) compressor isentropic efficiency (ηc),
b) compressor flow capacity (Γc),
c) gas generator turbine isentropic efficiency (ηgt),
d) gas generator turbine flow capacity (Γgt),
e) power turbine isentropic efficiency (ηpt),
f) power turbine flow capacity (Γ
Γpt).
A physical fault can be represented as a combination of one or more
listed component faults. To investigate the optimum set for different
physical faults, deteriorated performance has been simulated for ten
different component fault sets, which are listed in following Table 1.
A prior knowledge of performance parameter changes by fault set
is required to select dependent parameters for the M601 program.
Engine degradation affects fuel flow and power output, as a result
power output and fuel flow measurement could be selected to investigate the occurrence of failures. The selection of dependent parameters is needed to enable effective diagnostics. Effective diagnostics
would be achieved if selected dependent parameters are accurate,
easily measured, reliable and cheap.
Table 1 — Specification of fault sets
Table 2 — List of measured parameters
Neural network training and validation
A significant part of using a neural network is training. Training is
achieved through an iterative process where the training data is repeatedly fed into the network and it incrementally improves interconnection weights to match the network data to desired targets. However,
matching network outputs with desired targets do not say that the network has trained well. Therefore, testing a trained network with
a validation test is essential while a validation test does not include training data. The MATLAB Neural Network toolbox version 5 was
employed for this research to investigate the application of neural networks for diagnostics of the Walter M-601E turboprop engine. The
steps to develop a neural network for diagnostics of the M-601E engine are illustrated in Fig. 4. At first sight the diagnostics approach
appears to be very complicated. There are five networks in all there.
The diagnostics problem is very complicated and it is not possible use
only one neural network for this purpose.
NN1: Fault detection
Figure 5 describes the NN1 configuration as 8 inputs, 72 neurons in
the hidden layer and 1 output. The input layer receives input data
through the measurements. It is essential to have distinct input patterns, otherwise the network can not learn properly. Neurons in hidden layer were distributed in three sublayers: in first and in third sublayer were 35 neurons, in second only 2 neurons. It is well known
”bottle neck“ architecture usually used in diagnostics.
Figure 4 — Diagnostics approach by neural networks
13
L E T E C K Ý Z P R AV O D A J
T0
p0
nG
p 2C
T2C
-2
Two similar shapes of pattern with different target
results slow convergence training. The shape of pattern
here is referred to deviations from design point, for
example a shape of pattern is a rise in all of the parameters and one is a reduction in all the parameters. If
two patterns with similar shape but different outputs
are fed into the network it is more difficult for the network to be trained
b)
The network convergence rate was defined as error
changes during each Epoch. Since, all data types need
to achieve the same value of the error, the number of
epochs shows the mean convergence rate of the network. The Scaled Conjugate Gradient (SCG) algorithm
for training was used.
c)
The training time for each epoch will increase only if
the amount of data increases. This happens even if the
same data is repeated into the training set. The increase in time per epoch is approximately linear with number of data. It can be concluded that to have lower training time, feeding unnecessary data should be avoided.
Output layer
1 N
2
∑ (T − A)
N i =1
(4)
While N is the number of input patterns, T is the target value matrix
and A is the network output. The number of epochs increases until the
MSE reaches a minimum defined error. However, NNl training might
not achieve the desired minimum error of 1 * 10-4 due to the possibility of small rates of error change.
0
a)
Figure 5 — Developed Neural Network 1 configuration
MSE =
-0.5
on is taken for this level of degradation; the NN1 with associated configuration can detect faulty conditions with high reliability. The curve on Figure 6 is not Heavyside (or jump)
function because of Log-Sigmoid Transfer Function used in
output neuron of the neural network. The Hard-Limit Transfer
Function that would be expected here is not possible use in
case of Back-propagation learn algorithm.
To investigate the effects of input pattern on training performance, ten different sets of training data were provided and
fed to the NN1. All training was limited to similar termination
conditions.
Wfe
The training performance is represented by a number of epochs, performance gradient and Mean Square Error (MSE) which is defined by
Equation (4):
-1
Figure 6 — Network results nG degradation
P
Hidden layer
-1.5
Degradation [%]
T4C
Input layer
0.5
Network output
1.0
NN1: VALIDATION
0.0
The output values of the input nodes are modulated by the
connection weights, either being magnified if the connection
weight is positive and greater than 1.0, or being moderated if the
connection weight is between 0.0 and 1.0 and fed to each unit of
hidden layer. It was attempted to use one hidden layer with 72
associated nodes. However, if the training results are not satisfactory the number of nodes and hidden layer can be increased.
There is only one node in the output since desired output values
are set to ”0“ and ”1“; ”1“ shows a clean engine and ”0“ shows
a faulty engine, Training pattern included 16 462 sets of data for
degradations and atmospheric conditions variations. From the
total input pattern, 261 data are clean (design point & atmospheric conditions variations) and the others are degradation data
changing between [0.1 to 10.0] %. All input data are normalized
to a range of [-100% to 100%] by using design point conditions
from the performance model.
3/2007
Finally, including ambient temperature and barometric pressure changes into the training data can make training longer and
more complex, however the advantage is that the model will
not require a correction technique for inlet air changes.
NN2: Fault isolation
Table 3 — NN1 Training performance data
Figure 6 shows NN1 output estimation for gas generator speed
n G degradation at all nominal parameters. It was anticipated
that the network shows the value 0 for all outputs, but the
results show that network could not properly detect degradations for fault levels below 0.5%. Since no maintenance acti-
The network NN2 with the same architecture with NN1 was developed to isolate faulty components. There is only one node in output too
since desired output values are set to range [0.5 to 3.5]. The network
classifies failure into the following categories:
1 - Compressor Degradation. Failure is represented by output
ranges of (0.5, 1.5).
2 - Gas Generator Turbine Degradation. Failure is represented by
output range of (1.5, 2.5).
3 - Power Turbine Degradation. Failure is represented by output
range of (2.5, 3.5).
14
C Z E C H A E R O S PA C E P R O C E E D I N G S
Table 4 — NN2 Training performance data
The network training stopped when the minimum gradient criteria achieved. Table 4 summarizes the performance data of
the NN2 training.
Ten different sets of validation data were fed into NN2 to
verify the accuracy of network. The results are acceptable.
The NN2 could isolate compressor flow capacity degradations
above 0.5% from other degradations with 100% accuracy.
on-board thermocouples but problem is in their interconnection. It means it is not possible to reveal the burning-up of the
chamber cover in its initial stage.
What follows is a summary of the main conclusions attained
in this study:
1. Diagnostics precision is significantly dependent on the
chosen instrument sets. It was found that for a turboprop two shaft engine the following measurements
generate the best diagnostics accuracy: compressor exit
pressure, compressor exit temperature, fuel flow, inter
turbine temperature, output power and gas generator
spool speed. The optimum set determines the number
of input nodes for the neural network and used to provide training data.
2.
Different artificial neural networks have been examined for a variety of applications. For diagnostics, since
training data included target values, supervised algorithms were utilized. Consequently, Feed-Forward
Multi Layer Neural Network configuration with Backpropagation of Error was examined because of their
capability to learn by supervised algorithms.
3.
The Artificial Neural Networks learn from what they
feed into, as a result significant care should be taken to
prepare training data in respect of data uniformity and
format. The validation tests indicate that the Neural
Network has high reliability to detect and isolate failures for degradations more than 0.5% also it has acceptable accuracy compared to GPA techniques to quantify fault index.
4.
Training to detect two similar shapes of patterns with
different outputs is time-consuming. This was found by
detection of inlet air parameters changes from compressor efficiency degradations.
5.
The training data for this study included ambient temperature and barometric pressure variations. As
a result, training was longer and more complex; however after training no technique was required to correct
data for different ambient temperatures and barometric
pressures.
NN3 to NN5: Fault quantification
The final networks were designed to quantify the failure. It
could be used by a maintenance team to perform appropriate
action for such a level of faults. The networks were trained
using a supervised learning method. The SCG algorithms were
used to train networks.
Training data of NN3 included 693 patterns which are based
on the range of (223.15 to 323.15) K ambient temperature and
on the range of (25 to 115) kPa atmospheric pressure and (0.1
to 10.0) % compressor efficiency degradation. The results
from NN3 show that the neural network has acceptable accuracy for diagnostics in all range of degradations after training.
In addition the network could predict failure significantly faster than GPA.
The NN4 has in input layer only three nodes for ambient
temperature T 0 , barometric pressure p 0 and generator spool
speed n G . Network will calculate total inter turbine temperature T 4C . Gas turbine efficiency is subsequently given by comparing NN4 calculated temperature T 4C and by measured temperature T 4C . Network has the same architecture of the hidden
layer like previous, but training set is created by small number
of patterns. One node (neuron) is in output layer. Other details
for NN4 are presented in literature [3].
Architecture of network named NN5 is created: by three
nodes in input layer for ambient temperature T 0 , barometric
pressure p 0 and generator spool speed n G , followed by 72
nodes in hidden layer and by two nodes in output layer for torque on propeller shaft T k and for propeller shaft speed n P .
Training data set have to respect limitations for both output
parameters i.e. T k and n P .
Summary and conclusion
In this study a series of artificial neural networks were designed and trained to assess the ability of neural networks to diagnostics of gas turbine engines. The results indicate that neural networks are able to detect gas turbine failure, isolate faulty components and quantify fault index with satisfactory
accuracy and swiftness compared to GPA techniques. The
main benefits of this diagnostics approach are maintenance
time saving and capital costs reduction.
All the results presented in this study are derived for the
case of only one faulty part (compressor or gas generator turbine or power turbine. Combination of two (or three) faults is
not reflected.
It is a pity a fault of gas combustion chamber cannot be
detected. Inter turbine temperature T 4C is measured by nine
References:
[1]
British Standards Institution, 1984: Glossary or
Maintenance terms in Terotechnology; BS3811,
BSI, London
[2]
Kumar, Uday: Maintenance Management for
Modem Mining Equipment; Journal of Mines,
Metals & Fuels. Vol. 44. 1., 1996, pp. 25-29
[3]
Lamka, J.: Detekce diskredibility palubních termočlánků neuronovými sítěmi; Metodický postup,
Zpráva VZLÚ, 2007
15
L E T E C K Ý Z P R AV O D A J
3/2007
Evaluation Methodology of RTD Projects
Chapter — Verification of Evaluation
Methodology at Specific Projects
Hodnoticí metodika projektů výzkumu a vývoje — kapitola
Verifikace modelů na konkrétních projektech
Ing. Klára Grammetbauerová, Ph.D. / VZLÚ, Plc., Prague
Verification of suggested methodologies based on the identified RTD projects from aeronautics (ESA, NASA, DLR,
Darmstadt University, KTH - Royal Institute of Technology Sweden and TC AV ČR) is made. One of the methodologies
was identified as being used since April 2007 by the Government Council for Research and Development.
Na základě identifikovaných projektů V&V v letectví (projekty z ESA, NASA, DLR, z univerzity v Darmstadtu, KTH Royal Institute of Technology Švédsko a TC AV ČR) byla provedena verifikace variant hodnotících metodik. Byla také
zjištěna aplikace jedné z variant RVV.
Keywords: methodology, research & development projects, R&D, evaluation of.
1 Chapter: Verification of evaluation methodology
at specific projects
For introduced model options of project evaluation were checked different sources of specific projects. Research projects from aeronautics
were selected and a database was created of them. The total number
of project was 681 of which 21% (14 projects) were not suitable for
further testing within introduced options. 54 projects were subject to
testing. The main reason for excluding several projects was lack of
supporting information for the tested options.
Option 1 — Project as a target
The basic postulate of the R&D projects evaluation at the level of projects is the most objective statement describing how the targets of solved
project or other activity contribute to the target of research program and
its fulfillment. As a starting point it is necessary to understand the project target whether it fulfills the SMART2 condition of targets.
■
■
Table 1 — Tested projects according to source
■
■
■
Sources:
Airbus - http://www.airbus.com/
CIRC - http://www.circ.cz/dokums_raw/CIRC_CZ_both_sides.pdf
CRAFT - http://www.businessinfo.cz/cz/clanek/zdroje-financovani-z-eu/
DLR - http://www.dlr.de/
EUREKA - http://www.businessinfo.cz/clanek/zdroje-financovani-z-eu/
ESA - n/a
EU - http://ec.europa.eu/research/aeronautics/projects/
KTH - Royal Institute of Technology, http://researchprojects.kth.se/
MU Brno - http://ctt.muni.cz/toUTF8.cs/
Newspapers - http://www.businessinfo.cz/
NASA - http://www.nasa.gov/
PECS - http://www.czechspace.cz/
TC AV ČR - http://www.tc.cz/projekty/
TU Darmstadt - http://www.fsr.tu-darmstadt.de/
Four options were tested as typical research projects, with an important presumption of all approaches being to avoid the dangers of complex formulae and unwarranted assumptions.
■
■
■
■
Option 1 - Project as a target
Option 2 - Evaluation according to criteria - Project scoring
Option 3 - Project evaluation with the option of abandon the
project
Option 4 - Combination of the previous options
Offered solution has specific description.
Offered solution is measurable - it is stated how to recognize
when the solution is successful, solution verification should be
stated at the beginning.
The solution must meet the needs of recipient.
Offered solution has to be realistic.
It is stated the timeframe for being introduced into practical use.
This evaluating criterion (as identified in prior chapter of the evaluation methodology) was introduced in the similar spread by the RVV
(The Government Council for Research and Development) at the
negotiation related to the Government Evaluation Methodology on
April 18, 2007. This criterion should be used especially in the phase
of Decision making to which project should be used available financial sources. With respect to the general consensus at RVV it is not
necessary further testing of this option.
Option 2 — Evaluation according to criteria
— Project scoring
RTD evaluation should reflect effects of the research but should not be
done according to related activities. With respect to results — as the
most suitable approach was found Project Scoring as a result of existing
methodologies. It is evaluation according to group of criteria. There were
sum up several characteristics of successful projects to find some critical
points where it is possible to state the criteria. Afterwards, criteria were
identified that reflect how successful the project solution is.
In the group of tested projects the following characteristics were
found:
Table 2 — Typically stated items within tested projects
16
C Z E C H A E R O S PA C E P R O C E E D I N G S
According to the share of all of the items in the project percentage
is stated for each item in the total database of projects. There are also
some other characteristics supporting successful projects. Their
results are stated below.
Table 3 — General results of analysis
The European Union usually fund partially research projects; EU
support ranges between 52 - 81% while in average the share is 59 %
from the EU financial sources. Additional amount is financed either by
(i) the participant consortium, (ii) by government or (iii) by business
partner.
The average duration of the project is about 36 months. Only some
of the projects are evaluated according to the results (18%) and specifically stated implementation is in lower share (7%). Especially for
the RTD projects their implementation should be emphasized.
KTH - Royal Institute of Technology, http://researchprojects.kth.se/
MU Brno - http://ctt.muni.cz/toUTF8.cs/
Newspapers - http://www.businessinfo.cz/
NASA - http://www.nasa.gov/
PECS - http://www.czechspace.cz/
TC AV ČR - http://www.tc.cz/projekty/
TU Darmstadt - http://www.fsr.tu-darmstadt.de/
The risk might be decreased by using different sources of financing.
In the whole sample (see following table) there are two and more
sources of financing. The risk is shared as the investment is shared. I.e.
when 1/3 of the whole list investment projects is paid by one side and
2/3 by the other partner then the risk of abortion is in general divided
to 1/3 for the first partner and 2/3 for the second partner.
Following table shows further financial details of the projects.
Necessary numbers were available for 14 projects only. The column
of Contribution of the partners is in fact numerically identified risk and
also the price of the underlying instrument, i.e. the RTD project.
Table 4 — Financial details from the selected projects
Option 3 — Project evaluation with the option
to abandon the project
As a special case of evaluation was introduced project evaluation with
the option to abandon the project. Usually financial models can capture the essence of option value by directly incorporating existing
knowledge of uncertainty. The value of the option is derived from
the underlying instrument — the value of the financial amounts invested into the project. The calculation of the option value is included in
the initial costs of the projects and the value of the project also includes a risk3. The RTD projects evaluation in fact represents capital
budgeting of certain investment.
However, real options compare to financial options are not tradable (e.g. the solving company cannot sell the right to abandon the
research problem to another subject). Additionally, with real option
analysis, uncertainty inherent in investment projects is usually
accounted for by risk-adjusting probabilities (a technique known as
the equivalent martingale approach).
Chart 1 — RTD Project budget and EU share
The initial amount of investment depends on how the technology is
demanding. The most technology demanding projects identified here
are FLYSAFE — Airborne Integrated Systems for Safety Improvement, Flight Hazard Protection and All Weather Operations, SAFEE
Security of Aircraft in the Future European Environment and ANASTASIA — Airborne New and Advanced Satellite techniques and
Technologies in a System Integrated Approach. However, the share of
EU contribution is not the highest at the respective cases.
Option 4 — Combination of the previous options
As already stated the combination of option 2 and 3 uses the risk level
as a selecting criterion with respect to share of financing parties. The
risk level refers to the lose in case of unsuccessful project.
The following criteria for the scoring have been identified:
■
■
Sources:
Airbus - http://www.airbus.com/
CIRC - http://www.circ.cz/dokums_raw/CIRC_CZ_both_sides.pdf
CRAFT - http://www.businessinfo.cz/cz/clanek/zdroje-financovani-z-eu/
DLR - http://www.dlr.de/
EUREKA - http://www.businessinfo.cz/clanek/zdroje-financovani-z-eu/
ESA - n/a
EU - http://ec.europa.eu/research/aeronautics/projects/
■
■
■
■
■
Risk for the company specified in monetary units if project
being unsuccessful
Milestones of the project with clearly stated criteria of
success at each step
The objectives are clearly stated in the proposal
Implementation added
Time restriction criterion of the project
Existing combination of industrial and academic participants both types use slightly different financial sources
Specified number of participating research groups
17
■
■
■
L E T E C K Ý Z P R AV O D A J
Share in financing of the project and the total amount of
investment planned for the project. RTD projects should meet
the criteria as any other investment.
Finance - contribution of partners, mainly for risk layout statement.
Identified number of countries from where are the participating research groups - higher number might decrease volatility of the currency into the investment
Other tested aspects were found as non significant e.g. if is participating SME, if there is any proof of previous experience with this type
of projects, these criteria might be used as a check list.
2 Conclusion
Sixty eight projects were identified as a source of data for testing of
identified methodology options. The first option introduced in previous chapter was agreed by the government in the similar way within
the testing period Therefore other testing seems to be redundant. Option 2 — Evaluation according to criteria — project scoring brought
criteria that commonly occurred within the projects and also as general results from the database were found the following characteristics: clearly stated objectives, wider number of research group usually from larger amount of countries, the combination of industrial and
academic side might be advantage, the implementation is also added.
As an important criterion found was stated time restriction of the project together with identified milestones. In relatively smaller part, at
the aeronautical projects, was stated participation of SME.
Option 3 refers to project evaluation with the option to abandon the
project — it refers to limitation of the risk to abandon the project
according to the philosophy of the options.
As a reasonable result it is possible to take into account Option 4
— combination of the previous options that include the risk of the
whole project together with identified criteria within project scoring.
3 Attachment — Tested projects
Project ESA in cooperation with VZLÚ, a.s.
BUSINESS INFO http://www.businessinfo.cz/cz/clanek/zdroje-financovani-zeu/priklady-uspesnych-projektu-firem/1000522/38629/
EUREKA, http://www.businessinfo.cz/cz/clanek/zdroje-financovani-z-eu/priklady-uspesnych-projektu-firem/1000522/38629/
CRAFT - http://www.businessinfo.cz/cz/clanek/zdroje-financovani-z-eu/prikladyuspesnych-projektu-firem/1000522/38629/
TC AV ČR - http://www.strast.cz/projekty/
TC AV ČR - http://www.strast.cz/projekty/
TC AV ČR, http://www.strast.cz/projekty/
CIRC - -http://www.circ.cz/dokums_raw/CIRC_CZ_both_sides.pdf
TC AV ČR, http://www.locomotive-project.org/cms/
TC AV ČR, 6RP, www.omen-projekt.cz
PECS - http://www.czechspace.cz/en/science-and-research/cluster-ii
PECS - http://www.czechspace.cz/en/science-and-research/waves-and-turbulence
PECS - http://www.czechspace.cz/en/science-and-research/integral
PECS - http://soho.esa.int/science-e/www/area/index.cfm?fareaid=76
PECS - http://www.czechspace.cz/en/science-and-research/proba-2-dslp
PECS - http://www.czechspace.cz/en/science-and-research/proba-2-tpmu
PECS - http://www.czechspace.cz/en/science-and-research/vyvoj-software-prosisnet
DLR - http://www.dlr.de/as/en/desktopdefault.aspx/tabid-651/1102_read-1690/
DLR - http://www.dlr.de/as/desktopdefault.aspx/tabid-2035/2979_read-4582/
DLR http://www.dlr.de/as/Portaldata/5/Resources/dokumente/projekte/vela/The_VELA
_Project.pdf
DLR - http://www.dlr.de/as/en/desktopdefault.aspx/tabid-194/407_read-5453/
DLR - http://www.dlr.de/as/desktopdefault.aspx/tabid-3174/4820_read-6972/
DLR - http://www.dlr.de/as/en/desktopdefault.aspx/tabid-194/407_read-5434/
DLR - http://www.dlr.de/as/en/desktopdefault.aspx/tabid-194/407_read-5452/
DLR - http://www.dlr.de/as/en/desktopdefault.aspx/tabid-3384/5247_read-7664/
NASA - http://www.nasa.gov/centers/dryden/research/X45A/index.html
NASA - http://www.nasa.gov/centers/dryden/research/G-III/index.html
NASA - http://ipp.nasa.gov/ittp_success.htm
KTH - Royal Institute of Technology http://researchprojects.kth.se/index.php/kb_7928/io_9340/io.html?lastmodify=0&s
tr_search=&sortby=0&von=0&funding=0&thematic_area=0&thematic_area2=0
3/2007
&showregion=0&showfocus=0&showorga=0&str_showtype=&str_detail=&str_c
hange_type=&&&show_only_kb=0&show_details=1&show_all_persons=&sho
w_all_downloads=&show_all_funding=&show_all_crossreferences=&show_all_
keywords=&show_all_projects=&show_all_publications=&show_all_organisations=&show_all_furtherorgas=&show_all_funding=
KTH - Royal Institute of Technology http://researchprojects.kth.se/index.php/kb_7928/io_9619/io.html
KTH - http://researchprojects.kth.se/index.php/kb_7927/io_9088/io.html
KTHhttp://researchprojects.kth.se/index.php/kb_7928/io_9620/io.html?add_to_infobox=1&lastmodify=&str_search=&sortby=&von=0&funding=&thematic_area=
&thematic_area2=&showregion=&showfocus=&showorga=&str_showtype=&str
_detail=&str_change_type=&&&show_only_kb=&show_details=&show_all_per
sons=&show_all_downloads=&show_all_funding=&show_all_crossreferences=&show_all_keywords=&show_all_projects=&show_all_publications=&sho
w_all_organisations=&show_all_furtherorgas=&show_all_funding=&add_io=962
0
KTH http://researchprojects.kth.se/index.php/kb_7928/io_9241/io.html?add_to_infobox=1&lastmodify=&str_search=&sortby=&von=0&funding=&thematic_area=
&thematic_area2=&showregion=&showfocus=&showorga=&str_showtype=&str
_detail=&str_change_type=&&&show_only_kb=&show_details=&show_all_per
sons=&show_all_downloads=&show_all_funding=&show_all_crossreferences=&show_all_keywords=&show_all_projects=&show_all_publications=&sho
w_all_organisations=&show_all_furtherorgas=&show_all_funding=&add_io=924
1
EU - http://ec.europa.eu/research/aeronautics/projects/article_3666_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3720_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3698_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3701_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3702_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3703_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3700_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3699_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3705_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3704_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3712_en.html
EU - http://ec.europa.eu/research/transport/projects/article_3710_en.html
NASA - http://www.aoe.vt.edu/~mason/Mason_f/REVCON_NRA.pdf
EU http://ec.europa.eu/research/aeronautics/doc_pdf/project_synopses_vol1_en.pdf
EU http://ec.europa.eu/research/aeronautics/doc_pdf/project_synopses_vol1_en.pdf
EU http://ec.europa.eu/research/aeronautics/doc_pdf/project_synopses_vol1_en.pdf
Darmstadt - http://www.fsr.tu-darmstadt.de/research/projects/en_aixm.html
Darmstadt - http://www.fsr.tu-darmstadt.de/research/projects/en_isaware.html
Darmstadt - http://www.fsr.tu-darmstadt.de/research/projects/en_pilas.html
DLR - http://www.dlr.de/emma
DLR - http://www.dlr.de/emma2/
EU - http://www.x-at.co.uk/index.php?page=res_proj
EU - http://www.x-at.co.uk/pdf/Reccomp%20A1.pdf
NASA - http://www.nasa.gov/centers/dryden/research/HyperX/index.html
NASA - http://www1.dfrc.nasa.gov/gallery/photo/Fleet/HTML/EC73-3495.html
EU http://europa.eu/rapid/pressReleasesAction.do?reference=MEMO/07/243&format=HTML&aged=0&language=EN&guiLanguage=en
Airbus - http://www.airbus.com/en/worldwide/uk/research_development/composites_network.html
Airbus http://www.airbus.com/en/worldwide/uk/research_development/wing_research_pr
ogramme.html
NASA http://www.nasa.gov/mission_pages/constellation/ares/rocket_science.html
NASA - http://www.nasa.gov/centers/dryden/research/F-15B/index.html
NASA - http://www.nasa.gov/centers/dryden/history/pastprojects/X43/index.html
NASA - http://www.nasa.gov/centers/dryden/research/Altair/index.html
http://ctt.muni.cz/toUTF8.cs/
DLR - http://www.dlr.de/as/en/desktopdefault.aspx/tabid-3574/5578_read-8080/
DLR - http://www.dlr.de/as/en/desktopdefault.aspx/tabid-195/274_read-470/
Notes:
1
2
3
— as of September 2007
— SMART - Specific, Measurable, Aligned, Realistic, Timed
— The risk neutral approach opened the door to a host of option valuation
techniques that used binominal trees or the Monte Carlo method to
model future asset values.
18
C Z E C H A E R O S PA C E P R O C E E D I N G S
Analysis of Fatigue Crack Growth under the
Spectrum Loading
Výpočet šíření trhliny při zatěžování spektrem
Ing. Petr Augustin, Ph.D. / Institute of Aerospace Engineering, Brno University
of Technology
This article describes the methodology of numerical simulation of fatigue crack growth under the spectrum loading
and its application to integrally stiffened panels manufactured using high speed cutting. A load sequence generated
from the spectrum of Boeing 737 airplane similar to the standardized load sequence for flight simulation tests on
transport aircraft wing structures TWIST was applied. The methodology starts by the determination of stress
intensity factors from FEM results using crack closure technique. Subsequent crack growth analysis is done in
NASGRO and uses a description of crack growth rates by the Forman-Newman-de Koning relationship. Two crack
growth models were tested: non interaction and Willenborg model. Verification of analyses was based on tests of
integral panels and CCT specimens performed under the spectrum loading at the IAE laboratory. In order to
determine crack growth rate data, additional tests on CCT specimens using the constant amplitude loading were also
carried out.
V příspěvku je popsána metodika numerické simulace šíření únavové trhliny při zatěžování spektrem a její použití
v případě vyztužených integrálních panelů vyrobených vysokorychlostním obráběním. Je aplikována zatěžovací
sekvence vygenerovaná ze spektra letounu Boeing 737 se strukturou obdobou standardní zatěžovací sekvenci
dopravního letounu TWIST. Metodika simulace šíření trhliny je založena na stanovení součinitele intenzity napětí na
základě výsledků napěťové analýzy konstrukce metodou konečných prvků postupem označovaným jako crack closure
technique. Navazující výpočet šíření trhliny probíhá v programu NASGRO a používá popisu rychlosti šíření trhliny
rovnicí Forman-Newman-de Koning. Byly aplikovány dva modely šíření trhliny — model nezahrnující interakci kmitů
různých amplitud a Willenborgův model. Pro verifikaci výpočtů byly na Zkušebně letecké techniky leteckého ústavu
uskutečněny únavové zkoušky integrálních panelů a CCT vzorků při zatěžování spektrem. Pro stanovení materiálových
charakteristik byly navíc provedeny zkoušky na CCT vzorcích na jedné hladině zatížení.
Keywords: crack growth, spectrum loading, fatigue, fracture mechanics, integrally stiffened
panels.
Introduction
This paper deals with the topic of numerical simulation of fatigue
crack growth in the specific case of application of the spectrum (or
flight simulation) loading. It is linked to the study of crack propagation under the constant amplitude loading described previously in
[1]. Both investigations were done on integrally stiffened panels
with two stringers manufactured using high speed cutting. The
work was carried out within the scope of the 6th Framework Programme project DaToN - Innovative Fatigue and Damage Tolerance Methods for the Application of New Structural Concepts.
Description of load sequence
Relatively enough information concerning load spectra measured
on transport aircraft has been published to date. The following data
sources were considered in order to select the proper flight load
spectrum for the tests:
- Large database ONERA taken from reference [2] pertaining
to 1,781,548 flight hours flown by different aircraft operated
by British Airways.
- Database ACMS (Aircraft Condition Monitoring System) kept
at NLR [2]. The data were measured during 121,893 flight
hours representing 24,358 flights of Boeing 747 aircraft.
- Measurements on 17 Boeing 737-400 aircraft in the extent of
19,105 flight hours carried out by FAA within the Airborne
Data Monitoring Systems Research Project [3].
- MD82/83 data related to the same FAA project obtained from
7,120 flight hours [4].
- The data set published by FAA collected during 9,164 flight
hours of the Boeing 767-200ER aircraft operation [5].
- Results of the NASA VGH Program. The data were obtained
using the Digital Flight Recorder System of Boeing 727 aircraft during 1,765 flight hours [6].
- NASA VGH data of Boeing 747 aircraft collected during
1,689 flight hours [7].
- Spectrum of the standardized load sequence of transport aircraft wing structures TWIST [8].
Figure 1 provides comparison of flight load spectra considered. It
can be seen that all data with exception of the TWIST spectrum are
very similar. Substantial differences lie in the omission level defined during the extraction from acquired records and in the truncation of high loads related to an extent of the database. In case of the
ONERA spectrum, only peaks and valleys larger then ⏐Δn⏐ > 0.5
were recorded. It is obvious that an essential extrapolation in the
area of small loads would be necessary for a flight simulation fatigue test application. On the other hand, the data published in references 5 to 7 don't include the area of largest load factors. This is
the reason for selection of measurement performed on Boeing 737
airplane (reference [3]). This spectrum covers an adequate range of
load factors and in comparison to other spectra is also more severe. Ground load spectrum of B 737 aircraft in comparison with
other selected data is presented in Fig. 2.
An important question of the development of a load sequence is
the definition of the number of flights in the repetitive block defining the magnitude of the highest load included. In accordance
with many sources (for instance refs [8], [9]), the number of flights
was established as one tenth of the anticipated lifetime. The B 737
design service objective of 75,000 flights was obtained from [10].
It leads to the repetitive block of 7500 flights. Another decision that
19
L E T E C K Ý Z P R AV O D A J
Fig. 1 — Flight load spectra of transport airplanes (TWIST
exceedances per 1000 flights)
3/2007
Fig. 2 — Comparison of ground load spectra
Fig. 4 — Position of the most severe flights in the sequence and
magnitudes of largest load factors
appropriate load factor is the largest load factor selected from number of peaks pertaining to one flight randomly generated from the
ground load spectrum.
Methodology of simulation of crack growth
Fig. 3 — Examples of different flight types
has to be made is the omission of small loads. The spectrum of
flight loads normalized to 7500 flights was trimmed on the level
pertaining to cumulative frequency of 315 000. Provided that the
typical magnitude of 1g nominal flight stress is 70 MPa, the omission stress range is 12.46 MPa. This is a conservative choice in
comparison with the recommendations in references [9] and [11].
Continuous flight load spectrum was divided into ten discrete
levels and subsequently distributed into different flight types with
various severities. The technique of the definition of flight types in
the case of gust dominated spectra of a transport aircraft described
in reference [8] was adopted in this work. The main principle based
on the analysis of an influence of different weather conditions on
the gust load spectrum is that the spectra of different flight types
should have similar shape and the extreme value distribution of highest loads in flight types should approximate log-normal distribution. Consistently with the TWIST load sequence, the number of
flight types is the same as the number of load levels. Thus all flight
types denoted by letters from A to J can differ in terms of the magnitude of highest load level as is shown in Fig. 3.
The load sequence is created on a flight by flight basis and the
sequence of loads and flights is random. Generation of the load history is realized by the in-house computer program FLTSIM.
All flights are finished by insertion of one compression load. The
The approach starts with the calculation of stress intensity factor
function from finite element results obtained using
MSC.Patran/Nastran. Simulation of crack propagation in a real
structure requires determination of stress intensity factors for
a large number of crack configurations and that is why simple FE
models comprising the shell elements were built. The crack closure technique was adopted for calculation of stress intensity factor
values.
Subsequent analysis of crack propagation is done by the software package NASGRO [12]. All the analyses used the representation of crack growth rate data via the Forman-Newman-de Koning
equation. Two crack growth models were tested: non interaction
and Willenborg model. Simple non interaction model is fully based
on crack growth rates determined from constant amplitude tests. It
isn't able to reflect any retardation effects due to the load history
hence conservative predictions can be expected. Generalized Willenborg model deals with the crack growth retardation by implementation of the effective stress ratio
Reff =
K min − K R
K max − K R
(1)
that is used within the crack growth equation instead of the stress
ratio for the current applied cycle of loading R = Kmin / Kmax. Modified residual stress intensity
K R = φK RW
( 2)
20
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 6 — Crack growth rate data curve fit for 2024-T351
aluminium alloy
Fig. 5 — CCT specimen with antibuckling restraints used for constant
amplitude and spectrum tests at IAE laboratory
is defined using the Willenborg residual stress intensity factor
1
K
W
R
=K
OL
max
⎛
Δa ⎞ 2
⎜⎜1 −
⎟⎟ − K max
Z
OL ⎠
⎝
(3)
where KOL
max is the maximum stress intensity factor for the overload
cycle, Δa stands for the crack growth between the overload cycle
and the current cycle and ZOL is the size of the overload plastic
zone. The factor φ is defined as follows:
1−
φ=
ΔK th
K max − K min
( RSO − 1)
( 4)
where ΔK
Kth is the threshold stress intensity factor range and is the
shut-off value of the stress ratio KOL
max / Kmax. When this value is
exceeded, the crack growth is arrested.
A typical feature of integrally stiffened panels is branching of
cracks growing through the stiffener. Simulation of this phenomenon requires consideration of parallel propagation of two cracks
during the cycle-by-cycle computational procedure. This problem
was treated using the two-dimensional growth option in NASGRO.
Analysis of fatigue crack growth in the CCT
specimen
Before application to the DaToN panel, a calculation of crack
growth under the spectrum loading was performed for simple CCT
specimen geometry. Since an analytical solution of the stress intensity factor function is in this case known, it was possible to verify
the crack growth analysis done in NASGRO without an influence
of inaccuracy of stress intensity factor determination. Both the analyses and the flight simulation test used for verification (see Figure
Fig. 7 — Predicted crack growth in the CCT specimen under the flight
simulation loading versus result of the test (arrows indicate
application of the most severe flight A)
5) were performed on 97 mm wide and 2 mm thick CCT specimens.
The material and the thickness of coupons were identical with the
skin of the DaToN panels tested in Brno.
All the simulations of crack propagation described in this paper
are based on crack growth data composed of the data taken from
NAGRO database pertaining to stress ratio of R = 0.1 and the sheet
thickness up to 3 mm and the data for R = 0.5 obtained from constant amplitude tests on CCT specimens carried out at IAE laboratory. Figure 6 depicts the curve fit of these data.
Comparison of crack growth simulations with the test is shown
in Fig. 7. The best prediction was obtained for Generalized Willenborg model with the shut-off ratio set to 2.5 and the range pair
counting technique.
Crack growth in the DaToN panel under the
spectrum loading
The methodology described above was finally applied to the
DaToN two-stringer panel. The panel was made of 2024-T351 aluminium alloy using high speed cutting technique. Fatigue cracks in
the skin got started from the central saw cut with the size of 2a =
20 mm.
Half FE model of the panel is shown in Fig. 8. Load was uniformly distributed in the gripping area of the panel. Since the load
sequence includes compression loads, it was decided to perform the
testing with an antibending device in order to support the cracked
panel in compression. The antibending device consists of two plates placed in the central part on the panel. Deformed shape of the
21
L E T E C K Ý Z P R AV O D A J
3/2007
Fig. 9 — Stress intesity factor functions for skin crack
in the DaToN panel
Fig. 8 — Stress contour plot and deformed shape of the two-stringer
panel with antibending device
panel with the plates is depicted in Fig.8.
Stress intensity factor functions for cracks in the skin obtained
from nonlinear FEM analyses are shown in Fig. 9. It has appeared
that application of the antibending device leads to an acceleration of
crack growth in the stiffeners. It results in earlier complete failure
of the stiffeners indicated by a jump increase of the skin crack
stress intensity factor.
Prediction of crack propagation in the two-stringer panel under
the spectrum loading in comparison with the test carried out at IAE
laboratory is depicted in Fig. 10. The shut-off ratio value as well as
the counting technique was identical with the analysis of the CCT
specimen.
Conclusion
In this paper, the topics of a development of the load spectrum
and its application to the analysis of fatigue crack growth were
presented. Crack propagation in the simple CCT specimen as well
as in the two-stringer integral panel was investigated. Relatively
wide experimental program comprising both the constant ampli-
Fig. 10 — Prediction of crack growth in the skin of the DaToN HSC
panel under the spectrum loading in comparison with the test
tude and the spectrum tests on two analyzed geometries was
undertaken in order to acquire the crack growth rate data and
enable the verification of analyses.
Data 1978-1980: 1689 Hours; Report NASA CR181909, Vol. IV, 1989
References:
[1]
[2]
[3]
[4]
[5]
[6]
[7]
Augustin, P.: Prediction of Crack Growth in Integrally
Stiffened Panels; Czech Aerospace Proceedings, No.
2/2007, pp. 5-7
Jonge, J.B. de, Hol, P.A., Gelder, P.A. van: Reanalysis of
European Flight Loads Data; Report DOT/FAA/CT94/21, FAA, 1994
Rustenberg, J., Skinn, D., Tipps, D.O.: Statistical Loads
Data for Boeing 737-400 Aircraft in Commercial Operations; Report DOT/FAA/AR-98/28, FAA, 1998
Skinn, D., Tipps, D.O., Rustenberg, J., Statistical Loads
Data for MD-82/83 Aircraft in Commercial Operations;
Report DOT/FAA/AR-98/65, FAA, 1998
Tipps, D.O., Rustenberg, J., Skinn, D.: Statistical Loads
Data for Boeing 767-200ER Aircraft in Commercial Operations; Report DOT/FAA/AR-00/18, FAA, 2000
Crabill, N.L.: The NASA Digital VGH Program - Exploration of Methods and Final Results, Volume III - B727
Data 1978-1980: 1765 Hours; Report NASA CR181909, Vol. III, 1989
Crabill, N.L.: The NASA Digital VGH Program - Exploration of Methods and Final Results, Volume IV - B747
[8]
Jonge, J.B. de, Schütz, D., Lowak, H., Schijve, J.: A
Standardized Load Sequence for Flight Simulation Tests
on Transport Aircraft Wing Structures; Report NLR TR
73029U, 1973
[9]
Lanciotti, A., Lazzeri, L.: Effects of Spectrum Variations
on Fatigue Crack Growth; International Journal of Fatigue, Vol. 14, No.5 (1992), pp. 319-324
[10] Akdenitz, A.: The Impact of Mandated Aging Airplane
Programs on Jet Transport Airplane Scheduled Structural
Inspection Programs; Aircraft Engineering and Aerospace
Technology, Vol. 73, No. 1 (2001), pp. 4-15
[11] Abelkis, P.R.: Effect of Transport Aircraft Wing Loads
Spectrum Variation on Crack Growth. Effect of Load
Spectrum Variables on Fatigue Crack Initiation and Propagation; ASTM STP 714, D.F. Bryan and J.M. Potter,
Eds., American Society for Testing and Materials, 1980,
pp.143-169
[12] NASGRO Reference Manual, NASA Johnson Space
Center, Southwest Research Institute
22
C Z E C H A E R O S PA C E P R O C E E D I N G S
Measurement and Evaluation of the
Interior Noise in Aircraft
Měření a hodnocení vnitřního hluku v letadlech
Ing. Tomáš Salava, DrSc., Ing. Marcela Šloufová / VZLÚ, Plc., Prague
In 2006 an extended methodology of measuring and evaluation of the interior noise in airplanes has been proposed in
the framework of project A5 of the Aeronautical and Space Research Center (CLKV) at the Aeronautical Test and
Research Institute in Prague (VZLU). This paper first outlines the present state and basic philosophy behind the
newly proposed methodology. Next, examples of measurement results obtained in accordance with the proposed
extended methodology are presented and discussed.
Vnitřní hluk je významný faktor v hodnocení komfortu cestujících ale také v pracovních podmínkách posádky nebo
pilota v malém letadle. V roce 2006 byla ve VZLÚ v rámci úkolu A5 CLKV navržena rozšířená metodika měření a
hodnocení vnitřního hluku v letadlech. V tomto článku je nejprve naznačen současný stav a základní záměry a cíle nově
navrhované metodiky. Dále jsou prezentovány a komentovány příklady výsledků měření provedených podle nově
navrhované metodiky.
Keywords: Noise analysis, interior noise, noise abatement, propeller-driven aircraft.
Editorial Note: Some pictures to this article are printed in colour on
the inner back-cover page.
Introduction
Interior noise in aircraft is undoubtedly a significant factor both in
flight comfort of passengers, and in working environment of the crew.
Yet, it is not subject of airworthiness certification and of any specific
concern by ICAO. To our knowledge, no international regulations or
limits have been issued so far, relevant specifically to internal noise in
airplanes. The generic international and national health protection
noise regulations should, however, apply to passenger compartments
and crew or pilot cabins in aircraft, too.
As to the methods of measuring interior noise in aircraft, first specific international standard has been issued by ISO already in 1981.
Regrettably, activities in updating standardization of measuring methods for interior noise in aircraft have not been much sedulous last
years. As far as we know, the last issued and valid relevant standards
concerning interior noise measurement in aircraft are:
DIN ISO 5129 (2003): Acoustics — Measurement of sound pressure
levels in the interior of aircraft during flight
ISO 5129-2001: Acoustics — Measurement of sound pressure levels
in the interior of aircraft during flight
SAE ARP 1323A (1990): Type Measurements of Aircraft Interior
Sound Pressure Levels During Cruise
German standard DIN ISO 5129 is just a national equivalent of
ISO 5129-2001. SAE ARP 1323A is seventeen years old now. In
Czech Republic is still formally valid a Czech standard ČSN 31 0306
Acoustics -Measurement of the Interior Noise of Aircraft, which was
issued in 1983 and is the Czech equivalent of ISO 5129 (first edition)
from the year 1981 [1]. The last version of ISO 5129-2001 (third edition) [2] should, however, be taken for valid internationally and thus
also in Czech Republic.
However, ISO 5129 has been intended for large civil transport aircraft and it does not suit to small airplanes. Beside this, it implies some
other limitations, which restrict its wider application. In 2006 therefore an extended methodology for measurements and evaluation of
internal noise, especially in smaller and light airplanes, has been proposed in the framework of the project A5 of the Aeronautical and
Space Research Center (CLKV) at the Aeronautical Test and Research Institute in Prague (VZLÚ) [3].
In this paper there is first outlined the present state in measuring
and evaluation of interior noise in aircraft and the basic philosophy
behind the proposed extended methodology. Then, examples of measuring results, obtained accordingly and using also contemporary
advanced measuring equipment and software are presented and commented.
Measuring interior noise in aircraft according
to IEC 5129-2001
As it has been mentioned already, IEC 5129 has been intended for
standardization of measuring and reporting the interior noise levels in
large civil transport aircraft, namely in their different parts and at different places. For smaller airplanes only the basic requirements specified in this standard can be used accordingly. Beside this, IEC 51292001 requires measuring internal noise in aircraft only under the conditions of steady cruise flight, with no passengers and minimum crew.
As usual, IEC 5129-2001 starts with the normative references and
definitions of some basic terms. Next it specifies the general requirements on measuring instrumentation and on the testing and measuring
conditions. Overall sound pressure levels are to be reported in dBA
according to IEC 61672-1. Minimum required frequency range of the
measuring equipment and of the 1/3 octave spectrograms is 50 Hz to
10 kHz (in central frequencies of standard 1/3-octave filters). At least
16 second averaging time is required for reporting 1/3 octave spectra.
Prescribed are further the locations of the measuring microphones,
generally near the supposed ear positions of passengers and the crew.
Basic requirements are given also on microphone mounting. Typically for large aircraft, locations are divided to passenger compartments,
crew stations and sleeping quarters. Specified in detail are the aircraft
flight conditions and also basic configurations of aircraft interiors
during measurement. In detail specified are the data required to be
given in test reports.
New in IEC 5129-2001 is using recordings of noise samples for
further processing as standard practice. Minimum 30 second duration
of every recorded noise sample is required. Only very basic requirements are given on the recording facility. Also new in IEC 5129-2001
is introduction of the term ”uncertainty“ of measurements. An informative explanation with basic expression to calculate the expanded
uncertainties of measurements are in Annex A.
Noise regulation applicable to interiors of aircraft
To our knowledge, no international regulations and limits exist, relevant specifically to internal noise in airplanes. However, the generic
23
L E T E C K Ý Z P R AV O D A J
3/2007
Fig. 1 — Lightweight digital
recording set with a pocket size
precision digital recorder
and international health protection noise regulations should apply
fully also to noise in aircraft, as it has also been already mentioned.
Of basic relevance in this respect should be the standards concerning
noise exposure and noise-induced hearing impairment, namely ISO
1999-1990 Acoustics — Determination of occupational noise exposure and estimation of noise-induced hearing impairment [4].
Other relevant standards are e.g. IEC 60804.1985 or US standard
ANSI S1.25 1991 which concern mainly the requirements on measuring instrumentation and measuring the sound exposure levels. The
sound exposure (or noise doze) integrates basically the sound power
during the time when the noise is acting on a person [5]. Generic noise
exposure health limits are given e.g. in WHO Noise Exposure Criteria
[6]. Examples of the allowable noise exposure durations according to
[6] are in table 1.
Sound exposure is also dealt with in the EU Directive 2003/10/EC
”On the minimum health and safety requirements regarding the exposure of workers from physical agents“. Since 2006, this Directive is
implemented into the Order of Czech Government No. 148/2006
””About the health protection against the adverse impacts of noise and
vibration“. However, this directive too, contains no noise limits for
interior noise even in any transportation mean.
Next relevant standard valid in Czech Republic is ČSN ISO 19991993, which is the Czech equivalent of ISO 1999-1990 Acoustics —
Determination of occupational noise exposure and estimation of noiseinduced hearing impairment" [4]. According to this standard it is possible to evaluate the noise impact on the crew of an airplane from the
measurements complying with the requirements of IEC 5129-2001
(or still valid ČSN 31 0308-1983).
Till the year 2000, in Czech Republic, the noise limits for interior
noise in airplanes were determined in the Czech health protection
regulations as follows: LAmax= 80 dB for less than 2hour flight and 75
dB for longer flights. The contemporary health protection legislative
Fig. 3 — Recording of noise in a UL airplane — measuring microphone is clamped on the pilots headset, the small lightweight recorder
has pilot in his pocket
Fig. 2 — Calibration of the recording set using a precision calibrator
by Bruel & Kjaer
and regulations are presently subject to continuous changes, the detailed description of which would be far beyond the scope of this article.
Table 1 — Allowable exposure durations
1)
IEC 60804, no intermittent LC in excess of 140 dBC
Proposed extended methodics and measuring
instrumentation
Proposed extensions to ISO 5129-2001 aim mainly at interior noise in
small and light aircraft, down to ultra light airplanes. Beside this, in
all aircraft, measurement and evaluation interior noise is not limited
only to steady vertical cruise flight. Beside measurements aimed at
health protection and traveling comfort criteria, the extended methodology pursues also to the diagnostic analysis an noise abatement
applications.
Wide choice of noise measuring instruments and measuring systems is available now, which can be used for measurement and evaluation of noise in aircraft, beginning by low cost sound level meters
up to sophisticated ”high-end“ sound level meters-analyzers and computer assisted measuring systems. After years of practice with several
generations of precision sound level meters and analyzers mainly by
Bruel & Kjaer, we decided to concentrate on instrumentation based on
calibrated digital recordings and data logging.
For acquiring calibrated digital recordings of noise in airplanes, we
chose first the battery operated minidisk digital recorder-player
SONY MZ-NH700. This pocket sized device is equipped by very
accurate analog to digital converters and makes possible high precision digital recordings in non compressed 16 bit linear PCM format.
The only disadvantage of this otherwise excellent mini recorder-player is its use of magneto-optical recording on moving medium with rather high energy consumption.
Since 2007, low-cost semiprofessional pocket-size digital recorders
have been available, which use high capacity memory cards as recording medium. We tested such one produced by Edirol, type R-09. This
recorder makes possible 16 or 24bit linear PCM recordings with sampling frequency up to 48 kS/sec. It is shown in Fig. 1 together with the
miniature condensed microphone SONY ECM 610. In Fig. 2 is shown
calibration of this microphone and recorder set, using a precise calibrator by Bruel & Kjaer.
24
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 4 — Noise levels (dBA) in a cabin of an ultra light airplane during
take-off
Fig. 5 — Noise levels (dBA) in an ultra light airplane during a longer
flight (see text)
Frequency response of this recording set is 22 Hz - 18 kHz + 1,5
dB. With 16 bit quantization the theoretical dynamic range is 96 dB,
usable range is approx. 86 dB. With 24 bit quantization the theoretical dynamic range is 144 dB. Edirol R-09 in 24 bit quantization
mode provides the usable dynamic range definitely more than 100
dB. We found, however, the linear 16 bit quantization mode fully
satisfactory, with optimum sensitivity setting for sound pressure
levels range from 40 dB to 120 dB with 6 dB headroom.
An example of using this digital recording set in ultralight airplane is shown in Fig. 3. Measuring microphone is clamped to the
communication head set of the pilot. This instrumentation does not
bother the pilot at all. Beside applications in very small airplanes,
this set makes it possible to acquire highly accurate recordings of
noise in large aircraft under conditions of their regular flights and
without any attracting attention of passengers.
To ensure high accuracy, calibration recordings are added before and after any recording or flight sequence. Thus the overall
sound pressure level (SPL) uncertainity of the calibrated digital
recording is less then ( + ) 0,5 dB, plus the uncertainty of the used
precision certified calibrator, which is also less then 0,5 dB. Reachable overall sound pressure level uncertainity of our digital
recordings is therefore less then 1 dB. Recordings can be processed
immediately e.g. on a notebook or stored to be processed and archived later.
If a detailed analysis of the noise is not called for immediately
than lightweight, even low-cost data logging sound level meters
prove more advantageous then ”high-end“ ones, with many instantly working evaluation functions. Similarly, processing the recorded
or logged data on standard computers, using suitable software, proves more useful then using even extremely expensive, advanced
sound level meters-analyzers and expensive measuring systems.
Time recordings
Examples of measurements and measuring results
The output of ISO 5129-2001 are basically, ”one-number“ indicators
of noise levels, measured in different places of an airplane, however
only during its steady vertical cruise and without passengers. However, such description of the interior noise in airplanes is often inadequate, especially with small and light airplanes. More complete information mainly about the noise dependence on the flight regime and
about the noise properties are often needed. Following examples of
different measurements and measurement results according to the
proposed extended methodology should show more.
First five examples show that time recording of the noise in aircraft
give more useful information than a ”one-number“ indicator of the
noise in airplane. First example in Fig. 4 shows a record of the noise
levels in the cabin of an UL airplane during its take-off. The recording begins by a short idle period and motor test on the ground.
Then starts take-off run and after approx 1 minute begins climbing.
On the graph the noise levels are in dBA, and have been logged
every second. During the take-off the noise level has exceeded 100
dBA.
A longer recording of the noise levels in the cabin of the same
airplane is in Fig. 5. It was acquired during a sequence of loop
flights with simulated take-off for measurement of external noise
of the airplane. The recording starts again by a short idle and motor
test. Then follow take-off and seven overflights above the measuring point at full motor power. Maximum recorded noise level has
been 107 dBA this time.
Another example of a time record of the interior noise levels is
in Fig. 6. It was recorded in the cabin of a small, cargo, propeller,
two-engine airplane, during its ground tests. At the full power of
the engines the noise level in the pilot cabin has reached nearly 110
dBA ( and nearly 120 dBC ). In Fig. 6 are plotted sound levels in
dBA.
An example of noise levels in a large turbofan airplane is shown in
Fig. 7. The record was acquired during its take-off run followed by
fast climbing and climbing with reduced power of the engines.
Maximum recorded noise level during the take-off run exceeds 90
dBA, the mean level during climbing is 82 dBA.
In Fig. 8 there is a similar example of the interior noise recorded
this time in a regional category turbo-propeller, airplane. In this
graph captured is a longer time period including taxiing and longer
waiting for take-off clearance. The take-off starts at the time 15:18.
Maximum noise levels exceed also 90 dBA during the take-off run.
The mean noise level during the climbing at full power is about 88
dBA and decreases later below 80 dBA.
Spectral analysis
Sound pressure level, usually in dBA is most often used simple noise
”strength“ indicator. However, it does not say anything about the noise
properties. IEC 5129-2001 prescribes, beside the sound levels in dBA,
also measuring sound pressure levels over standard 1/3-octave filters.
The result of such measurement is 1/3-octave spectrum. In Fig. 9 is
1/3-octave spectrum of the noise sample from the same airplane as for
which the time record of the interior noise levels is shown in Fig. 8 and
25
Fig. 6 — Noise levels (dBA) in a small two-engine cargo airplane
during its ground tests
it holds for the full-power climb after take-off.
On the graph in Fig. 9, zero level on the scale corresponds to 100
dB sound pressure level. The graph describes the spectral properties
of the noise of turbo-propeller airplane during its fast climbing. Evidently dominant is here the sound level in the 1/3-octave band with
the central frequency 125 Hz. One-third octave data are used mainly
for computing the Effective Perceived Noise Level ( EPNdB ), which
is a more precise measure of the perceived noise ”strength“ or annoyance than are the noise levels in dBA.
For noise abatement diagnostics the 1/3-octave spectra are too
coarse. For this purpose high resolution spectra are used. In Fig. 10
there is high resolution spectrum of the same noise of which the 1/3
octave spectrum is in Fig. 9. It has been obtained by Fast Fourier
Transform (FFT) technique [8]. In this high-resolution spectrum there
are already well visible the dominant tonal components in the noise,
which look like very narrow, salient maxima. They are the harmonic
spectral components of the ”hum-like“ noise, generated by propellers.
Another example of high-resolution spectra is in Fig. 11, in this
case of the noise on a front seat of a large turbofan airplane, again
during climbing after take-off. In this spectrum distinct are the dense
high-frequency tonal components of the ”saw-like“ noise generated
by the turbofan blades. They are radiated from the propulsion unit forwardly. Finally, in Fig. 12 there is shown a typical FFT spectrum of
an ultra light airplane with a three-blade propeller and four-cylinder
four-stroke piston engine.
In Fig. 12, there are already demarked the main most prominent
harmonic spectral components, those which are generated by the propeller (Vn) and the ones, generates by the engine exhaust (Mn). At
full power climbing, the tonal component at approximately 100 Hz is
evidently dominant. Its level reaches 110 dB (lin). During take-off the
noise of the propeller is again evidently dominant. In this case the propeller and engine spectral components were discernible relatively
easily.
The airplane was equipped by three-blade propeller and four-cylinder four-tact engine with a reductor, having the reduction ratio 1:2,43.
The frequencies of the propeller and engine components must have
been in ratio 1:1,23. Frequency 100 Hz corresponds to the first, fundamental harmonics of the three-blade propeller at 2000 rpm. With
1:2,43 reduction ratio this corresponds to 4680 rpm of the engine. The
fundamental frequency of the exhaust noise under same conditions is
then 81,3 Hz.
Fig. 10 — High-resolution FFT spectrum of the same noise, of
which the 1/3-octave spectrum is in Fig. 9; narrow prominent maxi-
L E T E C K Ý Z P R AV O D A J
3/2007
Fig. 7 — Noise levels (dBA) at a front seat in the passenger compartment of a large turbofan airplane during take-off and fast climb
Fig. 8 — Noise levels (dBA) at a front seat in the passenger compartment of a regional turboprop aircraft; taxiing, longer waiting, take-off
at 13:17, followed by a 1,5 minute period of full-power fast climb and
continued climbing with reduced power
Fig. 9 — Result of the 1/3-octave analysis of the interior noise in
turbo-propeller airplane during fast climb after the take-off (in the
graph 0 dB=100dB SPL)
ma represent the components (harmonics) of the ”hum-like“ noise
generated by propellers
Spectrograms and more advanced evaluation
methods
In Figs. 9 to 12 were shown examples of short-time spectra, or shorttime integrated frequency spectra of interior noise in different airpla-
26
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 10 — High-resolution FFT spectrum of the same noise, of which
the 1/3-octave spectrum is in Fig. 9; narrow prominent maxima represent the components (harmonics) of the "hum-like" noise generated
by propellers
Fig. 11 — An example of a typical FFT spectrum of an interior noise at
a front seat in a large turbofan aircraft, during its fast climbing after
the take-off
nes. If the spectra change more in time, the three-dimensional spectrograms give more lucid information. In Fig. 13 there is an example
of a 3-D high-resolution FFT spectrogram of the interior noise in still
the same turbo-propeller airplane, this time during climbing with
decreased power of engines.
Another type of spectrogram is shown in Fig. 14. The third dimension, the sound level, is expressed usually by color scale. In Fig. 14
the originally color spectrogram is impaired by its conversion into
gray shades. This spectrogram shows the noise spectra in a turbo-propeller airplane during its take-off. Between fourth and five second the
engines go to full take-off power and the frequency of the first harmonic of the propeller noise increases from approx. 70 Hz to 120 Hz.
Traces of the first and several higher harmonics should be visible.
Still other 3-D graphic representations of spectrograms are used.
However, the ordinary spectrograms are basically successive sets of
short-time spectra. Higher resolution both in time and frequency
yields time-frequency distributions as e.g. the Wigner distribution [8].
Both the 3-D spectrograms and time-frequency distributions may look
very impressive with good graphics, but for many purposes further
processing is needed, mostly very time consuming if it is to be done
”by hand“.
F of the propeller noise decreases to nearly exact 100 Hz. Very similarly as F and the rotational speed of the propellers behave also the
sound levels Ltot, Lprop and Ldif.
The SPAD analysis shows in this case first of all, how much is the
propeller noise above all the other noise during the different flight
phases. However, this is not the only result of this particular SPAD
analysis. It shows e.g. also that decreasing propeller speed ( rpm ) by
a bit less then 20% results in approximately 12 dB decrease in propeller noise level, which means the decrease to 1/16 in sound power
radiated by propellers and is in very good agreement with the theory
of propeller noise generation.
SPAD analysis
SPAD analysis is based on so-called spectral decomposition (Spectral
Analysis and Decomposition) [9]. An example of graphical presentation of basic results of the SPAD analysis is in Fig. 15, obtained by
program SPAD ver. 0.7 [10]. Analyzed has been again the interior
noise of still the same turbo-propeller aircraft. The graph covers the
time sequence of 3 minutes which is demarked by a box in Fig. 8.
On the graph in Fig. 15 plotted are from top to bottom:
● Fundamental frequency of the propeller noise F [Hz]
● Overall sound pressure level Ltot [dBA]
● Sound pressure level of the extracted propeller noise components Lprop [dBA]
● Sound pressure level of the rest noise without the propeller
noise Lrst [dBA]
● Level difference Ldif = Lprop - Lrst
At the 10th second the engines go to full power. At 30th second the
take-off run ends and starts fast climbing. After some 90 seconds
climbing continues with reduced power of engines.
Let us follow first the fundamental frequency F of the propeller
noise. After the engines go to maximum take-off power and the airplane starts take-off run, F increases to 121 Hz and still a bit more
after the airplane starts climbing. After some 80 seconds the fast climbing ends and continues with reduced power. Fundamental frequency
Summary and conclusions
Interior noise in aircraft is significant factor both in flight comfort of
passengers, and in working environment of the crew. Though it is not
subject of airworthiness certification, the first international standard
for measuring noise in aircraft has been issued already in 1981. In
Czech Republic is still formally valid the Czech standard ČSN
31 0306 Acoustics-Measurement of the Interior Noise of Aircraft,
which was issued in 1983. It is the Czech equivalent of ISO 5129 from
the year 1981. The last version of ISO 5129-2001 (third edition) [2]
should, however, be taken for valid internationally and thus also in
Czech Republic.
A drawback of ISO 5129 is that it has been intended for large civil
transport aircraft. The output of ISO 5129-2001 are ”one-number“
indicators of noise levels in different places of an airplane, however
only during its steady vertical cruise flight and without passengers. In
2006 therefore an extended methodology for measurement and evaluation of internal noise, especially in smaller and light airplanes, has
been proposed in the framework of the project A5 of the Aeronautical and Space Research Center (CLKV) at the Aeronautical Test and
Research Institute in Prague (VZLÚ).
The proposed extensions to ISO 5129-2001 aim mainly at interior
noise in small and light airplanes, down to ultra light airplanes. Beside this, interior noise of all aircraft measurement and evaluation is not
limited to steady vertical cruise flight without passengers. Moreover,
the extended methodology pursues not only measurements aimed to
health protection and traveling comfort criteria, but also to the noise
abatement applications and noise diagnostics.
In this paper there is first outlined the present state and the basic
philosophy behind the proposed extended methodology. Then, examples of measurements done in accordance with the proposed extended
methodic are presented and discussed, together with using the contemporary measuring and recording technologies. Advanced low-cost
instrumentation for measuring and analysis of interior noise in airplanes were tested successfully and are recommended especially for testing smaller aircraft.
27
L E T E C K Ý Z P R AV O D A J
3/2007
The proposed extended methodology gives much more room for
advanced measuring, analysis and evaluation of interior noise in
aircrafts than gives ISO 5129-2001. It has proved especially well at
measuring internal noise in large aircraft under their regular flights
conditions and in testing smaller aircraft with even low-cost instrumentation.
Editorial Note: The remaining pictures to this article are
printed in colour on the inner back-cover page.
Fig. 12 — An example of the FFT spectrum of the noise in the cabin of
an ultra-light airplane during its climbing after the take-off; Vn — the
propeller components, Mn — engine (exhaust) components (for more
detail see text)
References:
[1]
ČSN 31 0306 Měření vnitřního hluku letadel (1983); ISO
5129-1981 Acoustics -Measurement of the Interior Noise of
Aircraft; (1981)
[2]
ISO 5129-2001 Acoustics -Measurement of the Interior
Noise of Aircraft; (third edition 2001)
[3]
http://www.euro.who.int/Noise/activities/20030123_3
[7]
Directive 2003/10/EC of the European Parliament and of
the Council on the minimum health and safety requirements regarding the exposure of workers to the risks arising from physical agents (noise) http://europa.eu.int/eurlex/en/dat/2003/l_042/l_04220030215en00380044.pdf
Šloufová M.: Vnitřní hluk letadel, měření a hodnocení,
metodika; VZLÚ Report R-4913 (2006)
[8]
R. G. Lyons: Understanding Digital Signal Processing;
Prentice Hall, 2004
[4]
ISO 1999-1990 Acoustics - Determination of occupational
noise exposure and estimation of noise-induced hearing
impairment (1990)
[9]
T. Salava, M. Šloufová: Spectral Decomposition in Noise
Abatement of Propeller Driven Airplanes; Czech Aerospace Proceedings, p. 22, No. 3, 2005
[5]
IEC 60804.1985 , US Standard ANSI S1.25 1991
[6]
WHO EU, Noise and Health Indicators;
[10] T. Salava: Program SPAD and its Use in Noise Abatement
of Propeller Airplanes; Czech Aerospace Proceedings, p.
14, No. 3, 2006
Composite Recycling — Technology,
Recycles, Their Parameters and Possible
Applications
Recyklace kompozitů — technologie, recykláty, jejich parametry a možné aplikace
Ing. Miroslav Valeš, Ing. Bedřich Štekner, Karel Cihelník / VZLÚ, Plc.,
Prague; Ing. Jan Grégr, Ing. Vladimír Kovačič / TU Liberec
The article deals with progress in solution of recycling of fibre composite with thermosetic matrix in the area of both
technology and methodology and also in the application of products resulted from composite waste treatment — recycles.
Článek je zaměřen na vývoj řešení problematiky recyklace vláknových kompozitů s termosetickou matricí, a to v oblasti
technologické, metodiické, a také v oblasti řešení využití produktů zpracování kompozitního odpadu - recyklátů.
Keywords: ecology, composites, plastics, thermosets, waste, recycling.
Introduction
One of characteristic features of the present industrial time is the effort
to use new types of material, which, in general, enable to improve
usage and technical characteristics in comparison to traditional constructional materials, especially the metallic ones. To give a few
examples — new materials can show better mechanical and physical
qualities as well as other qualities together with lower weight and, at
the same time, applications of these materials can enhance reliability,
safety and durability of a product. Such advanced materials, whether
in aeronautical or in other industries, are composites, especially composites with long-fibered reinforcement and different types of thermosetical matrix. However, utilization of composite materials is followed by an issue not solved yet - how to treat products made of these
materials after the expiration of their service life. The general trends
in waste disposal are minimum of simple disposal and on the opposite side strong emphasis on re-using of waste materials in production,
in brief recycling of these materials.
28
C Z E C H A E R O S PA C E P R O C E E D I N G S
The recycling of composite material has already been described in
other contributions [1], [2], nevertheless it is worth to remind, that this
issue has been investigated VZLÚ, Plc., for three years not only theoretically, but also practically. Currently, the processing of composite
waste is executed on an experimental equipment for thermal (pyrolysis) breakdown of thermosetic composite materials. This equipment
was designed and commissioned in 2006. The principle of the process
is based on controlled combustion of the composite matrix and separation of original reinforcing fibres. Products (recycles) obtained by
this procedure feature a possibility to be used in following manufacturing of new products.
°C T1
700
II.
III.
IV
°C T2
V.
0,21
°C T3
600
°C T4
0,2
% O2
500
ppm CO
0,19
400
ppm NO
ppm NO2
0,18
ppm SO2
300
ppm H2
0,17
°C TS
200
0,16
100
0,15
kg w eight
% CO2
0
0
50
100
150
200
0,14
300
250
exposition [m in]
quotient of gas [%]
temperature [oC] / quotient of gas [ppm]
I.
Fig. 2b
Fig. 2 — Decomposition phases of carbon-epoxy composite
Explanatory notes:
I.
II.
III.
IV.
V.
Sample: A
Size: 18.2040 mg
File: C:\TA\Data\TGA\Jana K\hasieiTGA\A.003
Operator: jk
Run Date: 20-Sep-2007 12:52
Instrument: TGA Q500 V3.1 Build 133
TGA
100
1.2
30.01min 94.11%
500
1.0
89.96min 81.83%
90
80
0.6
300
0.4
200
0.2
Research on thermal processing (pyrolysis)
70
In order to select optimal processing technology and obtain resultant
recycles featuring the most appropriate characteristics it is important
to know the used processing method in details. This investigation has
been performed from the very beginning of experimental work; in
result the basic phases of decomposition were established, primarily
on basis of weight decreases and local extremes of typical combustion products, which separates individual phases. The pictures below
show particular phases during processing of the carbon-epoxy composite by means of charts visualizing the real decomposition.
700
5250
325oC 440oC
II.
III.
600
545oC568oC
I
V
V.
°C T1
5240
°C T2
°C T3
500
5230
400
5220
300
5210
°C T4
% O2
ppm CO
% CO2
ppm NO
ppm NO2
505oC
200
ppm SO2
5200
467o
ppm H2
100
5190
0
5180
300
0
50
100
150
expos ition [m in]
200
250
weight [kg]
temperature [oC] / quotient of gas [ppm]
I.
Deriv. Weight (%/min)
0.8
400
Temperature (°C)
The technology of processing itself and using of it in waste disposal is
quite unique and in principle it enables to treat the major part of the
most common types of thermosetical composites. However, due to
the relatively short period of experimental use, this technology is currently verified and optimized in terms of experimental system hardware as well as processing technique — all of this in relation to obtaining recycles, in particular regenerated fibres, which feature as good
characteristics as possible. The associated issue is the possibility to
use products (recycles) obtained in this manner, which depends not
only on their characteristics, but also on successive steps of processing. These have been the main topics for research in 2007, especially in the area of carbon-epoxy materials.
The description of the process was elaborated in cooperation with
several external companies, such as Technical Institute of Fire Protection (given below as TUPO), where a number of analyses were
performed, for example fire tests of composite materials and screening chemical analysis of thermal decomposition of composite samples. A part of these analyses are tests by means of differential scanning calorimetry (DSC) and thermal analysis (TGA). Results of the
TGA from the TUPO Prague answer to findings in VZLÚ (see above).
The Figure 3 depicts the TGA of carbon-epoxy sample.
Weight (%)
Fig. 1 — Experimental equipment for thermal decomposition
(pyrolysis) of thermosetic composite materials
Zero decomposition sector
Primary decomposition sector
Secondary decomposition sector — the matrix is combusted
Latency sector
Degradation and burning off sector
°C TS
kg w eight
100
119.93min 65.17%
145.56min 62.57%
60
0
20
40
60
80
Time (min)
100
120
140
0.0
-0.2
160
Universal V4.1D TA Instruments
Fig. 3 — TGA of decomposition of carbon-epoxy sample
Technology optimization
Adjustments and optimization of the processing technology were
focused on refinement of the experimental system and on creation of
processing techniques for particular types of composite materials.
As for the experimental system it is necessary to point out, that the
system was designed so as to enable processing of composite material (waste) even without adjustments. However, there was a need for
some adjustments, in particular enhancement of data collection, corrections in cooling air flow, an alternative positioning of composite
inserted into thermal reactor and mainly, but not limited to, to minimize side effects of processing, this means to treat combustion gases
from the composite matrix. It should be noted, that even before these
adjustments the whole combustion process was monitored, in particular the content of typical products of combustion (e.g. COx, NOx, SOx
et. al.), by gas analyser Testo350XL and the flow of combustion products was controlled and diluted. But the laboratory tests and analyses, which were performed in TUPO Prague, should be considered as
29
Nevertheless, the hardware optimization of the technology described above is only one of solved areas. The same effort was devoted
to decomposition techniques as such, in particular the impact of the
processing character, minimum necessary and sufficient processing
temperatures, the duration time at such temperatures, the influence of
dimensions and shape of processed part (charge — the effect of
module M=V/S) and overall process procedure. The diagram below
shows some of the possible cases of the procedure, which have been
optimized in relation to entry parameters.
Process ingMap (carbon-e poxy)
600
temperature [oC]
500
t = fce( tl. )
Process D
400
Process C
Process B3
300
Process B2
Process B1
Process A
200
Process 0
100
0
0
15
30
45
60
75
90
105
120
135
150
165
180
195
210
225
240
255
270
285
300
315
time [min.]
Figure 7 - some of the typical procedures
Tab.1
T
b 1 Application
A li i off procedures
d
given
i
in
i Fi
Fig. 7
Another example is practical verification of sufficient time stability for
matrix combustion, which depends on thickness of the processed
composite part.
TimeMap (carbon-epoxy)
7
6
5
thickness [mm]
more precise and complex; these tests focused on chemical analyses
of concentrations of typical constituents in products of combustion
from decomposition of different types of composites in inert or noninert atmosphere and on determination of total toxicity.
The processing of three types of samples (carbon-epoxy, carbonbismaleimide and glass-phenol) was tested in non-inert atmosphere
and one sample (carbon-epoxy) was tested in inert atmosphere. By
comparing the result of this experiment it could be found, that the carbon-epoxy samples processed in inert nitrogen show the lowest toxicity, on the other side, the highest toxicity were found during processing in normal non-inert atmosphere for the carbon-bismaleimide
material while lower values of toxicity were detected for carbonepoxy material and glass-phenol material respectively.
Considering the above mentioned facts it could be stated, that
completion of the experimental system by active filtering of the
combustion products could be consider a very important part of the
project. The original filter system consisted of mechanical, two stages
stainless steel filter capable to catch hard particles in flowing gas
medium. This proved to be insufficient from the point of view of
elimination of combustion products; therefore the whole system was
completed by relatively mighty filters based on active coal. The new
filter unit contains 16 filter cartridges filled with 32 kg of the
Silcarbon SC40 active coal, in shape of granules of 3-4 mm in
diameter and 4-8 mm in length. This active coal is a carbon product
used for cleaning gases and air from harmful pollutants and for
reducing odours. It is characterised by very fine pores and large
internal surface — c. 1,100 m2.g-1, due to which this active coal can
adsorbe a wide spectrum of substances (substances are bound to the
internal surface).
The active coal filter system was inserted into the experimental
system at the displacement part and sized for gas flow c. 2,160 m3.h-1.
Saturation of the filter cartridges is monitored on the base of weight
increase and by determination of pressure difference between the
filter input and output using couple of pressure sensors — U-pipes.
Even the first tests with filter unit installed showed considerable
advance in project solution and in minimization of undesirable
products of processing.
3/2007
L E T E C K Ý Z P R AV O D A J
4
3
2
1
0
0
10
20
30
good
40
50
60
70
80
90
tim e [m in]
bad
Figure 8 - Experimental verification of dividing line between sufficient
and insufficient time stability
Recycles properties
Figures 4, 5, and 6 — Filter system
based on active coal, pressure difference indicator and filter cartridge
filled with active coal
The integral part of recycling issue should be a determination of resulting characteristics of products obtained — the recycles, in correlation with the type of processed material and technological parameters
of procedure. It is beyond all disputes that each using of recycled
material is in close correlation with material characteristics. So the aim
is to find such an optimization of the process method, which will result
in product featuring the best characteristics possible, or, in case of
composite materials with fibre reinforcement, regenerated fibres with
minimum degradation of properties.
This particular problem were worked out in close co-operation with
the Czech Technical University in Prague, Faculty of Mechanical
Engineering, Department of Manufacturing Technology and mainly,
but not limited to, with the Technical University of Liberec, Department of Textile Materials and Department of Chemistry, where different experiments and analyses were executed with regenerated fibres
prepared in various processing conditions. For example, one of these
30
C Z E C H A E R O S PA C E P R O C E E D I N G S
tests was analyses of fibre degradation (decreases determined optically or gravimetrically), or changes in mechanical properties, which
were determined on tension test device TIRAtest2300 (e.g. strength,
modulus, elongation…). Some of the results are showed in following
tables and pictures. The different levels of damage of regenerated fibres are illustrated on the pictures taken from the scanning electron
microscope TESCAN VEGA TS 5130. There is clearly visible
a distinct increase in fibre degradation due to higher temperatures
during processing in non-inert atmosphere, where material decreases,
fibre burning out or fibrillation are occurring.
The diameters of carbon fibres
8,43
The diameters of the mono-fibres
[ m]
9
8,26
7,71
8
7,43
7,38
7,07
6,39
7
5,55
6
4,94 5,06
5
4
3
2
1
0
starting
550°C - N2
550°C - air
The technology type of processing
600°C - air
w eight analysis
650°C - air
optical microscopy
Figure 9 — The influence of processing conditions on loss of carbon
fibres
SEM image of virgin carbon fibre
The material decreases of carbon fibres
70
64
Decrease [%]
57,3
56,7
60
50
40
29,7
28,6
30
20
4,45
10
0
5,37
4
0
0
starting
550°C N2
The technological type of the processing
550°C air
600°C air
weight decrease [%]
650°C air
volume decrease [%]
Figure 10 — The influence of processing conditions on loss of carbon
fibres
Rm [MPa] / E [GPa] / A
[%]
The mechanical properties of carbon fibres
1731
1496
91 1,53
105 2,04
1122 102
1,89
550°C - N2
550°C - air
The technological type of processing
600°C - air
Rm [MPa]
820 104
1,52
650°C - air
E [GPa]
ductility [%]
Figure 11 — The influence of processing conditions on mechanical
properties of carbon fibres
The possible applications
The aim of all above mentioned activities is to develop and verify such
methods and solutions, which will enable re-usage of composite
waste in any form — i.e. recycling. Focusing on products resulting
from procedures mentioned hereinbefore, the regenerated fibres could
be considered as the most important recycles. There are many possibilities, how to use the regenerated fibres, for example as raw material for new materials made by the powder metallurgy, in paper technology, in general as a short fibre reinforcement in different types of
matrix, filling material, in SMC, BMC, DMC, 3D textiles, as reinforced barrier coating systems and other.
In view of progress in this project we have started to deal with this
problem this year in several ways. One of them is the application of
regenerated fibres as a filling agent in thermoplastics materials, this
means in the products made by traditional plastic technologies. The
contribution of filling for example by carbon fibres is in enhancement
Two SEM images of carbon fibre by +600° C on air
31
L E T E C K Ý Z P R AV O D A J
3/2007
SEM image of carbon fibre by +550° C in nitrogen
SEM image of carbon fibre by +550° C on air
Three SEM images of carbon fibre by +650° C on air
of mechanical properties (for example toughness, stiffness) as well as
in possibility to ensure the electrostatic discharge.
The usage of regenerated fibres in mentioned types of particle
composites is possible on condition of fibre preliminary treatment,
especially milling to suitable length. At the beginning this preliminary treatment caused big troubles, as the regenerated fibres are very
soft, with diameter about 4-6 μm and arranged as a very disparate
material without any orientation and with high level of infiltration between particular layers and fibres. Due to this it is difficult to secure
the suitable orientation of fibres in relation to blades. The next problem encountered during the experiments was the length of fibres.
Experiments proved, that it is desirable to achieve mean length of the
cut fibres no more than 0.1 mm due to the fact, that longer fibres
agglomerates and their processibility grows worse.
The filling of the cut fibres as such into the thermoplastic matrix
was executed in co-operation with the Tomas Bata University in Zlin,
Polymer Centre, where the first samples of material filled by recycled
fibres from VZLÚ were produced and where the first rheological tests
were executed. The values of the specific electrical conductivity of filled and non-filled material were determined. As expected, material
filled by as little as 15 % of carbon fibres shows conductivity of 0.1
32
C Z E C H A E R O S PA C E P R O C E E D I N G S
S/m, which represents quite good value. Even though the results from
the early experiments are known, the whole project is at the beginning
and that's why it would be before the right time to come to any conclusion. However it is obvious that this is a promising way of using the
recycled carbon fibres. The important aspect of usage will be also the
economic parameters of recycles, especially regenerated fibres.
The pictures below depict the sample of common PP+SEBS material (10 %) and the sample of the same material, which is stuffed by
15 % recycled and cutted carbon fibre.
Conclusion
The recycling of fibre reinforced composite materials with thermosetic matrix is a vast issue, which includes the processing technology
itself for various charges, the creation of recycle with specific characteristics and consecutive treatment for different applications in practice. Currently, the focus of the project is being shifted from the processing techniques to application of recycles and it is possible to state,
that this is the area with a significant potential for development of new
materials utilizing regenerated fibres (recycles). Some of the first
results of this work are presented in this contribution.
It is necessary to point out, that all this issue can be seen from various aspects — not only the technical one, but also economic as well as
Figure
g
13 —The sample
p of PP+SEBS material and the sample
p of the
same material stuffed by carbon fibre
riveting, with improved quality. This new joining method is being proenvironmental aspects. In spite of considerable progress in composite recycling there are still a lot of unknown and, consequently, there
is still a lot of work to do in the future.
Literature:
[1]
Valeš, M., Kachlík, P.: Introduction to Problems of Thermosetic
Composite Materials Recycling; Czech Aerospace Proceedings, No.
3, 2005
[2]
Valeš, M., Štekner, B.: Experimental Test System for Fibrous Thermosetting Composites Breakdown; Czech Aerospace Proceedings,
No. 3, 2006
New FSW Equipment at VZLÚ,
Ú Plc.
Nové vybavení pro třecí svařování ve VZLÚ, a.s.
Ing. Petr Bělský / VZLÚ, Plc., Prague
The article gives short information about latest activities of VZLÚ, Plc. in area of friction stir welding. Within the
frame of the R&D works in the project ARC-B2 and participation in EU project LOSTIR a new monitoring and
clamping equipment was developed. It ensures high quality friction stir welding on conventional and NC milling
machines. The paper summarizes results of the tests of the new equipment. Finally, short information about next
planned research FSW activities in VZLÚ,Plc. is presented.
Příspěvek informuje o současných aktivitách VZLÚ, a.s. v oblasti třecího svařování. V rámci výzkumných prací na
projektu CLKV-B2 a účasti na evropském projektu LOSTIR bylo vyvinuto nové monitorovací a upínací zařízení, které
výrazným způsobem zkvalitňují frikční svařování na konvenčních i NC frézovacích strojích. Článek shrnuje výsledky
zkoušek tohoto nového vybavení a informuje o plánovaných aktivitách v nejbližším období.
Keywords: friction stir welding, monitoring, quality, clamping fixture.
1. Introduction
As an innovative joining technology for lightweight metal alloys,
Friction Stir Welding (FSW) attracts more and more interests from
manufacturing industries. Due to its numerous process benefits, such
as high product efficiency, perfect mechanical property, lower structure distortion and automatic machinery process, friction stir welding
is often depicted as a revolutionary welding method in the new century. Since it was invented by The Welding Institute (TWI-Cambridge,
UK) in 1991, through over decade of research and development, friction stir welding has become technically mature and is gradually applied into various industrial manufacturing fields, such as ship-building,
train-car manufacturing, aerospace, power engineering, civil construction and so on.
In the worldwide aircraft manufacturing industry, even as a prevailing trend in joining of lightweight metal alloy structures, FSW found
wide utilization too. Traditional technology of riveting fails to satisfy
recent enhanced demands for production progressive low-cost airframe structures. Riveting is simple and very well mastered joining method but it is very time-consuming technology. There are other disadvantages too. Friction Stir Welding makes possible joining speeds 6
times faster than automated riveting or 60 times faster than manual
ved to have the great potential to be engaged in the manufacturing of
the aircraft wing structures, wing boxes, airframes, fuselages, tailored
blanks of doors and windows, cargo assembly structures and so on.
The substantial progresses were made for example by Eclipse Aviation, Boeing and Airbus.
VZLÚ, Plc. officially started its own FSW research activities in
September 2006. The first experiments were focused mainly on tests
of basic principles and general demonstrations of the modern joining
method. Standard manual knee-type milling machine and simple
clamping fixture enabling realization of very short linear welds (c.
180 mm) were used. This equipment was not able to measure and
control all necessary welding parameters, and, as such, the FSW joint
quality produced by this system cannot be assured. That is why there
was made a decision to use new NC milling machine equipped with
a monitoring system, modern FSW tools and a new special clamping
fixture enabling realization of long linear and 2D curved welds.
This paper describes features and capabilities of the new equipment and summarizes first results of its tests in VZLU, Plc.
2. Welding machine
At the present time special FSW machines are commercially available
from several manufacturers licensed by TWI Ltd. (ESAB AB, CRAW-
33
L E T E C K Ý Z P R AV O D A J
3/2007
annular rotating antenna, with the information being transmitted to
a static calliper style coupling module. In addition there is a separate
housed electronics module that processes the information into a suitable format for interfacing with a PC.
The information gathered by the Lowstir device is displayed to the
operator in a clear and straightforward manner using a laptop PC running Labview. The instrument panel displays real time numerical values of forces, torque, the temperature adjacent to the system electronics
and (if desired) the tool temperature. The system also has the capability to add real-time event markers to allow correlation between process
conditions/stages and the recorded data. The main display screen has
buttons to start and stop recording of data. Alternatively an automatic
trigger facility exists for initiating the recording of data. The display also
shows the current captured data values for the weld in progress indicating whether they are within the acceptable range for satisfactory welding. The display also has a multi-graph facility where the user can
select which sensor values are displayed.
4. New FSW Tools
Fig. 1 — Monitoring system LOWSTIR
FORD-SWIFT, HITACHI LTD, MTS SYSTEMS… etc.). Unfortunately these purpose built machines are generally very expensive (above
250 000 Euros) and thus unsuitable for preliminary R&D activities of
VZLÚ, Plc. That is why a low-cost solution of the problem was sought.
Finally an adaptation of standard milling machine revealed as ideal
because the machine can be used for both machining in the workshop
and R&D FSW activities. The main ”FSW machine“ of VZLÚ, Plc. is
new NC bed-type milling machine FSG 80 A2 from TOS Kuřim. This
machine has large table clamping surface 2000x800 mm, high main
spindle motor output — 19.3 kW (spindle speed range: 20-4000 rpm,
max.torque: 1000 Nm) and control system HEIDENHAIN-iTNC530.
Adaptation of the machine for FSW technology required some small
design modifications. Most considerable changes were connected with
increasing maximal feed trust in axes ”Z“ from 15 to 20 kN. This important parameter affects maximal thickness of welded materials. Milling
machines also lacked any process monitoring capabilities required to
ensure high quality FSW joints. That is why it was equipped with special monitoring system LOWSTIR (see below).
3. Monitoring system
VZLÚ-ARC participated in EU project LOSTIR. The main objective of
the project was development of a low cost torque/force monitoring
device for conversion milling machines to FSW. VZLÚ, Plc. is the first
user of the system in the world. The weld monitoring system has been
developed to accurately measure the vertical and horizontal forces and
torque on the tool. The sensor is machined from one piece of high grade
stainless steel, heat treated for maximum strength and stability. The
sensor design allows for various taper sizes to be attached to accommodate the requirements of the user. The data gathered can be directly
linked to the acceptance or otherwise of the weld. In addition, the device has the capability to monitor two user defined temperatures via thermocouples, one to be attached to the FSW tool, the second will act as
a safety cut out to protect the integral telemetry circuit, monitoring the
temperature at the interface between the tool holder and the weld monitoring system. All of the power requirements for the sensor are transmitted using wireless digital telemetry. This also supports two-way
communication for calibration and data collection. The sensor has an
The welding parameters and tool geometry play a major role in deciding the weld quality. Progresses of FSW technology achieved in the
world are above all the consequence of the new FSW tool development
and a better process understanding. Nevertheless, one of the most
important challenges is still the design of the tool shoulder-pin system
to assure a good quality weld and to reduce the loads during the process. First FSW trials in VZLÚ, Plc. were performed with very simple
tools. No threads or scrolls on pin and shoulder were used. The tools
had concave shoulder and cylindrical rounded pin (see Fig. 2).
New FSW tools from Suffolk Precision Aerospace Ltd. have special scrolled shoulder and threaded pin with three flutes. The tool has
been designed to operate within the target range of torques, forces and
processing rates that can be achieved by standard milling machines.
The force exerted during the welding process is influenced by the
Fig. 2 — a) old concave FSW tool; b) new scroll type FSW tool
dimensions and profile of the tool shoulder. The FSW tool has a profiled shoulder to permit welding at 0° tilt, negating the need for a specialised head or tilting table. Body of the tools (shoulder) is made of
H13 tool steel and pin is made of high-strength (8G) inbus screw.
5. Vacuum clamping fixture
Sufficient clamping of the weldments during welding is one of the key
requirements for realization of superior FSW welds. First preliminary
experiments were carried out in year 2006 on a small various-purpose
fixture with maximum length of welds up to c. 180 mm. Unfortunately
this parameter was very limiting because it did not make possible production of longer fatigue specimens or evaluation of quality variation
for long welds. That is why in frame of R&D activities in the project
ARC-B2 a new special clamping fixture combining vacuum system
Horst-Witte and standard mechanical elements was developed. This
new design concept enables realization of long welds (up to 1200 mm)
and clamping of wide range of materials (incl. non-magnetic) and thicknesses. Applied vacuum system consists of a mobile modular unit Witte
82150 with suction capacity 63 m3/h, manifold vacuum distributor and
8 pcs. of modular grid type chucks 300x200x32,5 mm with grid
34
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 3 — Vacuum clamping fixture
Fig. 5 — Typical record obtained with LOWSTIR system during measuring of force response to welding across dividing line
12,5 mm. The type of chucks is generally used for clamping simple
geometrically shaped workpieces (weldments) with or without cutouts.
Special seals laid out according to the weldment contour (but incorporating the vacuum supply bore) help to even out pronounced uneven or
rough surfaces, whilst still securely holding the weldment during the
application. The fixture itself is divided into four clamping areas that can
be selected individually, thus allowing maximum flexibility with regard
to the size of the weldment.
6. Tests of the new FSW equipment
In the first instance the monitoring system LOWSTIR was tested on
manual milling machine FGU 40 (11 kW / 2000 Nm / 35-1800rpm) at
lower rotation speed. These experiments were focused on examination
of general possibilities of the apparatus, sensitivity measurement and
cooling capability. Overall 18 butt welds of 3mm thick sheets AA7075
were realized. Influence of welding parameters on forces acting on
FSW tool during welding was studied. Rotation speed was varied in the
range 280-1120 rpm and welding speed in the range 56-315 mm/min).
Welds No. 1-4 were carried out at ”higher“ rotation speed (1120 rpm)
and with shoulder diameter ∅15 mm. These two parameters caused
excessive heat generation and considerable distortions. That is why
shoulder diameter was reduced to ?12 mm and rotational speed decreased (Welds No. 5-18). These adjustments are limited because low rotational speed cause increasing penetration resistance during plunge period and weld force during welding. Excessive welding speed and very
low rotation speed can be the cause of fracture of the pin (see Fig. 4).
Next activities were also focused on testing of sensitivity of the
LOWSTIR system. The tests were based on force response measuring
during welding across dividing line between two 3 mm thick aluminium
sheets (AA7075) with different sizes of gaps (0, 0.1, 0.2 mm). The welding process was carried out with the tool rotating at 710 rpm and at
feed rates of 56 or 224 mm/min, with a 1,5° tilt angle and a 0.15 mm
plunge. Typical record of the experiment for gap 0.1 mm and feed rate
224 mm/min is shown in the Fig. 5. All performed tests confirmed sufficient sensitivity of measuring DOWN FORCE and TOOL TORQUE.
Tests of the new FSW tools with scrolled shoulder confirmed fact
that the tools is possible to use with tilt angle 0° (perpendicular to the
surface). It is very good finding because VZLÚ, Plc. plans to use 3 axis
NC milling machine and these tools will allow welding along 2D curve
(not only straight welds). The FSW tools worked very well but the
weak part of the tools was pin. Two pins were broken because of crosssection attenuation with flutes and insufficient strength. New pins made
of maraging steel or MP159 will be used for the next FSW activities.
Unfortunately manufacturing of the new vacuum-clamping fixture is
delayed. That is why tests of the equipment will not be performed until
November 2007.
7. Planned FSW activities in VZLÚ, Plc.
a) Confrontation of FSW technology with other modern welding
methods
Most advantages of friction stir welding result from the fact that the
process takes place in the solid phase below the melting point of the
materials to be joined. It results in lower heat load of weldments and
thereby low distortions, shrinkage and excellent mechanical properties.
New fusion welding methods (CMT, laser, electron beam) offer
also similar advantages and simultaneously high productivity. Planned R&D activities will be focused on confrontation these joining
methods.
b) Optimalization of welding parameters for lap joints of FSW
demonstrator
First static and fatigue tests of lap FSW joints were already performed
in VZLÚ, Plc. In the next year these activities will continue with the
aid of the new equipment. Above-mentioned vacuum clamping fixture was designed for realization of the demonstrator representative part
of airframe structure (section of wing).
Acknowledgements
These R&D activities were carried out in the ARC supported by the
Czech Ministry of Education, Youth and Sports. The author would
like to thank consortium partners of EU project LOSTIR as follows;
Peter Lewis of Applied Measurements Ltd UK; Peter Cheetham, Sigmapi Systems Ltd UK; Andy Ezeilo of TWI Ltd UK; Olga Mishina
of Sapa Technology Sweden; Andrew Wescott of BAE Systems UK;
Adam Pietras of Instytut Spawalnictwa Poland, Asun Rivero of Fatronik System Spain and Paul Wilson of Suffolk Precision Ltd UK.
References:
Fig. 4 — FSW tool failure indication
[1]
Kallee S. W., Nicholas E. D. and Thomas W. M.: 'Friction stir welding — invention, innovations and applications'; Tube International,
Vol. 21, No. 117, July-August 2001
[2]
Bělský P.: 'Preliminary Tests of Friction Stir Welding'; Czech Aerospace Proceedings, No. 3, pp. 24-26, VZLÚ, Plc., Prague, 2006
[3]
Bělský P.: 'Friction Stir Welding of Aircraft Structures'; Czech Aerospace Proceedings No. 3, pp. 15-17, VZLÚ, Plc., Prague, 2003
35
L E T E C K Ý Z P R AV O D A J
3/2007
Buckling
kl
of Shells
h ll
Stabilitní úlohy skořepin
Ing. Karel Patočka; Ing.
Ing
Ing Tomáš Jamróz; Jiří Had / VZLÚ,
VZLÚ Plc.,
Plc Prague
In the article the authors describe experimental results of buckling tests due to shear loading of plates.
plates These results
are compared to nonlinear calculations conducted by FEA.
V článku jsou popsány výsledky z experimentu při zkouškách ztráty stability při smykovém zatěžování desek. Tyto
výsledky jsou srovnány s nelineárními výpočty provedenými při MKP výpočtech.
Keywords: nonlinear buckling, composite plate.
1.1 Introduction
You can describe the response of a structure to load as deformation
energy spent or work done. In case of loading the structure comes
through a series of states and goes along a certain energy curve. But
if it suddenly begins to follow another energy curve, this is named
structure instability or buckling. One of a large number of stability
tasks is buckling of shells or thin-walled structures, respectively.
1.2 Solution to static buckling
In addition to analytical solutions to stability problems, since recently
there has also been a numerical apparatus that especially in concert
with the FEA methods brings a powerful tool to solve general tasks
from mechanics of flexible bodies. In direct numerical approach there
are still lots of complications.
As time goes on several methods have been developed. A basic one
and the oldest of them from historical point of view is a so called linear elastic stability method. Its principle lies in a resolution of eigenvalue problem of a structure and follows with determination of critical
load. This procedure is advantageous especially due to simple model
preparation and fast solution. But linear elastic buckling (LEB) otherwise called as linear prebuckling (LPB) has significant limitations that
give this procedure sense rather like initial estimation for thin-walled
structures.
Basic limitation of LEB are the following:
I. Conservative loading, thus only static problems are considered
II. One of the main limitation that is supposed in this solution is so
called symmetri
y
ical bifurcatiion, i.e. the only cases to be considered. If it is evident from intitial estimation, that one branch in
energy solution is more likely to occur then the other one the
calculation using LEB gives no reliable sense.
III. Deformation should be small by break down.
IV. Assumption of elastic behaviour up to buckling point.
V. Forces are not dependent on displacement. In LEB theory are
not so called ”follower forces“.
VI. LEB isn't appropriate method for structures, that are sensitive
to imperfection, i.e. small deflection from ideal geometric
shape. Such behaviour is shown in Fig.1.
Finding eigenvalues of a structure you solve basically the equation
(KM + λKG1) z = 0
where KM is material and KG is geometric part of tangent stiffness
matrix z are eigenvectors and λ is multiplicative constant describing
critical force.
Using nonlinear methods is more advanced approach how to describe buckling behaviour. The latter provides possibility to involve both
field of geometric nonlinearitis (large displacement, …) and material
nonlinearities (plasticity, …). So you are able to simulate postbuckling
phenomenons. Major differences between theoretical model and real
part is an existence of many perturbations like load excentricity, local
material inhomogenity, shape deflections (on the other hand model is
mostly characterized by ideal geometry) etc. Thin-walled structures
show inclination to imperfections and it is known, that the imperfections significantly cut resulting load capacity. There is a possibility how
get the model closer to the real part. Introducing these imperfection
into the numerical model is the way. You can virtually do it by applying additional loads to the structure forcing it to behave appropriately
in direction of corresponding state of the smallest potencial energy
needed. Alternative method is to enforce such behaviour by specifying dicplacement on particular structure nodes.
The magnitude and shape of displacement depends on specific setting. The most exact method is to establish displacement on account
of measured real imperfections and this way you can take into consideration the technological influences or pertiently other possible ones.
If such data aren't available it is possible to estimate the shape of
imperfections. In case of plates you can set imperfections like first
eigenshape from linear buckling analysis, thereby enforcing behaviour
of plate through the first shape by estimation that this shape corresponds to the value of minimum potencial energy. However, in structures highly sensitive to imperfections (e.g . cylindrical shells), the
first eigenshape need not correspond to minimum of potencial energy
and therefore imperfections need to be set in shape closer to the real
deflection from ideal geometry, thus for instance in shape of second
eigenshape or possibly in combination.
Determination of critical force itself is intimately bound up with the
numerical process. Basic approach in calculation is by dint of Newton
iterative method, when structure is step by step loaded by force divided into individual increments and the process lasts as long as the tangent stiffness matrix is positive. Then the calculation stops. Force corresponding to this state is critical buckling force. Other possibility is
a load through the use of force controlled by arg-lenght procedure as
far as to the postbuckling region. Force value, that matches sudden
displacement growth, is the critical force.
1.3 Experimental buckling measurements
Fig. 1 - Differences in
linear and nonlinear
buckling
Plates clamped in a frame were measured to obtain experimental data.
The frame is created so that induced shear in plates, which were carbon fiber reinforced composite made of eight prepregs of thickness
36
C Z E C H A E R O S PA C E P R O C E E D I N G S
of 1.36 mm. Orientations of layup was the following: [45]8, [-45]8,
[0]8, [0/90]2S, [90/0]2S, [45/-45]2S, [-45/45]2S.
Fig. 2a - Dimensions of frame
Fig. 2b - Clamping of composite
F
plate
1.3.1 Mathematical model
Both model of frame and model of plate were created in MSC.Patran
for numerical FEA calculations. There were used shell elements
QUAD4. Connection between them was simulated by multipoint constraints RBE2. Calculations were in MSC.MARC. As mentioned
above initial imperfections were introduced by virtue of linear buckling analysis like first eigenshape. Then the nonlinear calculations
were carried out. Load was set in 3 blocks (subcases). In first subcase initial force was divided equally corresponding to linear elastic
area. Second subcase was concentrated on the region of buckling.
Both in the latter and the third subcase the load was controled byl the
value of incremental force according to arc-length procedure (Riks
Ramm), which makes it possible to calculate so called snap through
or snap back problems.
Fig. 4 — Fig. 4 scaled for different imperfections
1.3.2 Isotropic plate
First an analysis of isotropic material was carried out to check model
and its sensitivity to the magnitude of initial imperfection. The shape
of imperfection was again like the first eigenshape in linear analysis.
The plate was made of duralumin sheet i.e. according to specification
of ČSN 42 4203.63 standard, its thickness was 1.0mm.
In figures you can see the influence of imperfection especially on
the initial buckling curve, while influence on stiffness is small.
Fig. 5 — Corner point displacement for different imperfections in FEA
Fig
and experiment
1.3.3 Composite plate
As in the case of isotropic material initial imperfection was introduced
due to first eigenshape. Nonlinear calculations were made as well.
Fig. 3 — Point displacement in the middle of isotropic plate
Fig. 6 — First eigenshape of plate (layup [45/-45]2S)
37
L E T E C K Ý Z P R AV O D A J
3/2007
Fig. 7 — Deformation of composite plate (layup [45/-45]2S)
Fig. 8 — Plate just before
failure [45/-45]2S)
Fig. 9 — Loading courses of composite plates
1.4 Conclusion
In Fig. 7 you can see displacement of plate whose shapes correspond
well to real displacement. The model shows also fair views of the
spots with the biggest stress thereby providing visualization of supposed failure point.
Nevertheless there are some other mechanisms, that are not involved in and lead to substantial resulting stiffness of the model as compared to reality.
Boundary conditions play a great role especially in the joints between plate and frame. Joints are realized like absolutely stiff in the
model, which does not agree with reality. Another model deviation
against experiment creates fibers failure mechanism in area of holes
for screws.
As to buckling point in isotropic material, you can see large plate
buckling in series of figures (Figs. 4, 5 and 6). It starts at a force of
about 3000N. Local buckling at a force of 600N is apparently due to
plate unevenness, but it is not the global buckling yet.
Regarding composite plates the conditions are more complicated
because the plates have of a order higher stiffness and displacements
are not so large. Deformation in places of largest deformation ar by
buckling at the same order with the plate thickness as well as the
dependence force vs. displacement of a point in the corner does not
show large curvature. Since no displacements were measured in a surface of the plate the estimation of critical buckling point is difficult and
model is more complex and more complicated as well.
In Fig. 10 it is shown courses of force vs. displacement in composite panels with different layup orientation. It is clear that except from
final region that belongs to the damage, there are no significant curvature changes and hence no significant indication of buckling.
One of practicable methods for critical buckling point determination on the basis of FEM calculations with respect to composite
plates was mentioned in the article. It is suitable to use simplier
case during debugging of mathematical model. For this purpose
we used experimental data of isotropic material. The model sensitivity to values of initial imperfection was also tested for the isotropic material. The initial deflection shape wasn't modified for
the sake of absence of experimental data in surface of plate.
It was shown that existence of imperfection is necessary for
nonlinear stability calculation. Its parameters affect especially
initial behaviour of plate.
It is misleading to consider unique magnitude of buckling load
with respect to composite plate. Unlike large displacements of
isotropic materials the deflections of composite materials are not
so considerable and after buckling the plate bears considerable
load capacity before total failure occurred.
Mathematical model provides good insight into resulting displacement and peak stress spots, nevertheless achieved deflections are smaller due to distinctively stiff behaviour of plate The
composite model has to be evidently refined on the contrary to
isotropic model, which relate very good with experiment.
References:
[1]
Featherston C.A.: Imperfection sensitivity of flat plates
under combined compression and shear; Elsevier Science,
2000
[2]
MSC.Marc Volume C: Program input
[3]
MSC.Marc Volume A: Theory and User information
[4]
MSC.Nastran: Handbook for nonlinear analysis
[5]
Zpráva grantového projektu Tandem T7
[6]
Schneider M. H., Halcomb J. R.: Stability analysis of perfect and imperfect cylinders using MSC/Nastran linear
and nonlinear buckling
38
C Z E C H A E R O S PA C E P R O C E E D I N G S
Numerical Study of Steady and Unsteady
Flow in a Centrifugal Compressor
Porovnání stacionárního a nestacionárního řešení
odstředivého kompresoru
Ing. Jan Tůma / VZLÚ, Plc., Prague
The article
i compares results off stationary
i
and non-stationary
i
numerical
i analysis
i off flow
f
conditions
ii
in
i centrifugal
if
compressor. Object of the comparison is to identify differences between the two methods. Second point is to consider
stationary analysis in terms of optimizing process applications.
Porovnání stacionární a nestacionárního přístupu z hlediska použití pro následnou optimalizaci proudění v radiálního
kompresoru. Srovnání odlišností řešení proudových podmínek použitím obou metod.
Keywords: Centrifugal compressor, Non-stationary, Blade.
Introduction
Boundary Condition
The article relates to the following research task: numerical modeling
of the flow in rotating devices. It compares accuracy of unsteady and
steady flow simulation in radial compressor (Fig. 1). Large variety of
problems originates in the unsteadiness of the flow. This unsteadiness
is created by relative motion in the geometry, by flow instabilities and
by multi component interactions. Large scale non-stationary simulations currently consume very large amounts of CPU time. The goal of
our project is to identify and solve problems with using stationary
state approach in optimization fluid passage process. The results could
also be used for other turbomachinery cases.
Two kinds of interfaces were used to solve the interaction between
rotor and stator:
Mixing Plane and Sliding Mash. Spalart-Allmaras was used as
a turbulence model. The solver was ”density based“ and ”implicit“
with second order accuracy. The Spalart-Allmaras model is a relatively simple one-equation model. It combines relatively short time of
computing and good results in displaying the development of boundary layers. It is gaining popularity in turbomachinery applications.
The fluid was assumed to be air, modeled as an ideal gas.
Following boundary conditions were used in computation: pressure inlet and pressure outlet or outlet vent (choking the outlet). Pressure outlet has been used as a standard condition in all manner of problems solved by computational fluid dynamics. As for stability of
computation, this condition proved to be very demanding. Outlet choking (outlet vent) was more stable and more compatible with the reality of the machine.
Flow field in impeller
The analysis of the flow conditions in impeller compressor was performed in regime of back pressure corresponding close to the design
point i.e. for back pressure close to the bottom limit of working area.
The Figures 5 and 6 illustrate flow field (relative path line and contour of relative or absolute Mach number) in impeller and diffuser for
non-stationary calculation.
Fig. 1 — Pathline through impeller and diffuser solved by CFX with
showed grid
Computational geometry and grid
The focus here will be on the High-Speed Centrifugal Compressor,
which is computationally more demanding. It consists of a radial-axial
impeller, vaneless and vane centrifugal diffuser and an axial diffuser.
For numeric simulation half machine was used. The final hexahedral
mash consisted of approximately one million cells.
Compressor grid consisted of O-grid around the blade and the
H-type grid in the passage. The part of computational geometry you
can see on Figure 1. Computation grid was generated using the preprocessor Gambit and the numerical solution was found with the help
of the CFD FLUENT.
The computing grids had between three and one million cells but
there was limit of solver license count and therefore was chosen coarse grid. But it is worked on implementation rotating models to other
solver which is not limited by license.
Fig. 2 — Controlling plane in compressor
12345-
rotational plane, touching leading edge of impeller blades
cylindrical plane, close of trailing edge of impeller blades (Discharge surface)
cylindrical plane, touching leading edge of diffuser blades
cylindrical plane, touching trailing edge of diffuser blades
rotational plane, touching trailing edge of diffuser blades
39
L E T E C K Ý Z P R AV O D A J
3/2007
The flow is most intensive on the pressure side of blades and splitters (as shown in Figure 5). Pathlines which begin on impeller inlet
plane get to low speed cores only sporadically, because the cores are
formed by flow from gap tip around leading edge area.
mean value of angle in tangential direction,
⎛c ⎞
α = arctan ⎜⎜ m ⎟⎟
⎝ ct ⎠
efficiency evaluation from total state with correction
interface influence
Stationary method of interface model ”Mixing Plane“
NonStationary method of interface model ”Slidingmash“
Boundary condition; chock the outlet (Outlet Vent)
Boundary condition; pressure outlet (Pressure outlet)
isentropic efficiency from total state with correction
interface influence,
Fig. 4 — Load of blade and
d splitter
litt 50% h
height
i ht off the blade,
non-stationary solution
Design constrains were the cause of a much higher value of Mach number than is recommended (especially on leading edge tip of impeller
blade) for our high-speed case. The compressor becomes unstable when
the operating line crosses the surge line. This is the so-called pinch point
in the stability map. The calculated performances were placed in Tab. 1.
The flow is first computed for steady-state condition (using mixing
plane concept) for half of the machine. Results of the mixing plane simulations were used to initialize non-stationary simulations. Comparison of
the load of the blades is shown in Figs. 2 and 3. Due to adverse pressure gradient, the compressor flow is prone to flow separation and a proper modeling of the turbulence in the near wall region is mandatory.
Final integral mean values from controlling planes (Fig. 2) concerning
both methods are in the Tab. 1.
Numerical simulation by both aforesaid methods shows that flow field
is comparable in both cases. Stationary method displays all essential areas
of loss similarly to the non-stationary one (see Figure. 3 and 4: load of the
blades). This is true in design regime of the machine. Simulation in an area
close to surge line failed to be stable and to be reach convergence.
Strong instabilities also occurred in cases of diminishing mass flow
which results in a change of the attack angle (as follows from velocity
triangles). This is similar to the area of choking. In this case as well, stationary methods did not display flow conditions of the machine rightly.
η=
Conclusions
Tab.1 — Integral mean value computed in one working point
Notation:
T, Tt —
Mass-Weighted Average mean value of total and static Temperature
Ts, Tt,s —
Mass-Weighted Average mean value of flow past
isentropic compression on appropriate value of static,
κ −1
total pressure from total flow state on inlet
⎛ p ⎞κ
⎟
⎜
T
T
=
Tt,0 = 288,15 K, pt,0 = 101325 Pa,
s
t,0 ⎜
⎟
⎝ p t,0 ⎠
α—
ημ —
MP, ST —
SM, NS OV —
P—
η—
ΔTt ,s
ΔTt
Load of the blades
It could be said that the stationary method (Mixing Plane) can be used
on the design point and operating conditions on a speed line. It means the
area from choke to near stall. Strong instability occurred out of this area
and therefore it is necessary to used non-stationary method ”Sliding
Mash“. The same is valid for the interaction between rotor and stator. It
should be emphasized that the accuracy of solution grows by using
unnon-stationary method and on the contrary the stationary method is
a reasonable approximation as long as there is no significant reverse flow
through mixing plane interface. If reverse flow occurs, stationary solution will not reflect real flow.
Editorial Note: The remaining pictures mentioned in this
article are printed in colour on the inner back-cover page.
References:
Fig. 3 — Load of blade and splitter 50% height of the blade,
stationary solution
Results
The solution was performed for one speed line to examine the flow
through the compressor. Computing of this compressor was shown quite
demanding as for solution stability and it was highly time consuming.
[1]
Hečl, P.: Analýza proudění ve dvoustupňovém jednohřídelovém
odstředivém kompresoru; 2005. 20 s.
[2]
Eckard, D.: Detail flow investigation within a high-speed centrifugal
compressor impeller; Trans. ASME, 1976
[3]
Eckard, D.: Flow analysis of radial and backswept centrifugal compressor impeller part I: Flow measurements using a laser velocimeter. In performance prediction of centrifugal pumps and compressors; ASME, 1980
[4]
FLUENT 6.3 — Manual Fluent Inc.
40
C Z E C H A E R O S PA C E P R O C E E D I N G S
Industrial Measurements of Frequency
Characteristics of Small Sport Aircraft
Průmyslová měření frekvenčních charakteristik malých sportovních
letounů
Ing. Karel Weigel, Ing. Tomáš Kostroun, Doc. Ing. Svatomír Slavík, CSc.
/ Department of Aerospace, Czech Technical University in Prague
The paper describes a system for ground frequency tests of LSA-category
LSA category aircraft.
aircraft Described are basic results of tests
on aircraft which were made in the year 2007.
Charakteristika měřicího zařízení. Hardwarová a softwarová vybavenost Metodika měření. Způsoby vyhodnocování
vlastních tvarů a frekvencí. Vývojová a provozní specifika.Přehled frekvenčních zkoušek provedených na malých
sportovních letounech výrobců ČR. Ukázky výsledky měření a jejich vyhodnocení. Záměry a cíle dalšího vývoje měřicí
soustavy a zpracování výsledků.
Keywords: Ground Vibration Test, GVT, Aeroelasticity, Flutter, Frequency Tests.
Introduction
A system for ground frequency tests was developed in cooperation
with TL-electronic company. The system is designed for tests of microlight, ultralight and LSA category of aircraft.
Hardware and software components
The system consists of an 11channel analyzer, a two-channnel signal
generator, power amplifiers and electrodynamic exciters. Additional
components are used for connecting sensors and exciters on the structure and device for lifting the aircraft. The dynamic signal analyzer is
compatible with piezoelectric accelerometers of IEPE standards. Sensors of 14g and 70g range are used for measuring.
Measured data are post processed in ME scope software. This software from time domain signal data calculates natural frequencies by
FFT and natural modes shapes by frequency response function. More
dynamic data functions as autocorrelation or coherence are available
too.
Measuring procedure
Natural frequencies and shapes of the main structure are determined
on soft suspended aircraft. The observed parts - wing or rear fuselage
with tails surfaces - are excited by electrodynamic exciter situated
under the lifting surfaces. The exciter is connected to the structure by
vacuum cup. The measuring of the rudders run with aircraft placed on
the ground. The observed parameters are main frequencies of the
whole control system.
The structure is always excited by sweep sine function. Used frequencies are from 2 to 100 Hz. Time domain data of output (forces)
and input (accelerations) are processed by FFT, after them the Frequency Response Functions are calculated.
The capability of the logger allows using 2 channels for output and
8 channels for input. The measuring processed on a pre-prepared set
of point. The points are situated by twos where former is on leading
edge and the latter is on trailing edge. Typically, each point is measured in two axes — direct and vertical. The number of points is usually about eighty on the drag and twenty on the rudders. For each
weight configuration it is necessary to take app. Eleven measuring
cycles on the wing, 8 measuring cycles on the fuselage and tail. Rudders take app. 6 measuring cycles that are no depending on aircraft
weight. The testing time is, depending on number of weight configuration, 5 to 8 days.
Fig. 1 — Set of the measuring points
Fig. 2 — Exciting of T tail of Allegro
Determination of the natural frequencies
and their shapes
Transformed frequency domain data of FRF's in each point are imported
to the ME scope software. Natural frequencies are selected according to
amplitude values in Real and Imaginary component and Phase shift.
41
L E T E C K Ý Z P R AV O D A J
3/2007
Fig. 3 — Animation of modal shapes in
ME scope software
Mode shapes are calculated from values of FRF complex vector. The
output data are arranged to tables optimized for aeroelastic software.
The aircraft
Since end of 2006 when the system was completed till October 2007
four frequency tests were carried out. The aircrafts were various
designs from classical high-wing monoplane to the faster light sport
aircraft. The aircraft has been tested: Atec Faeta, Schempp-Hirt M-7
Ornis, Aveko VL-3 and Fantasy Air Allegro.
Fig. 6 — Aveko VL-3
The Aveko VL-3 is a two-seat high-performance low-wing ultralight
airplane. The plane has been made with optimal use of advanced
materials such as polyurethane and carbon composites. The fuselage
of the plane consists of composites sandwich shells where the main
construction element is carbon fabric.
Fig. 4 — Atec Faeta
The Atec Faeta is a two-seat low-wing ultralight aircraft with a fixed
tricycle undercarriage. It is an advanced low-wing side-by-side twoseater, built in mixed construction of a carbon composites and wood.
Fig. 7 — Fantasy Air Allegro
The last one Fantasy Air Allegro is a single-engine two-seat (side-byside) braced high-wing airplane.
Illustrative results of measuring
Fig. 5 — Schempp-Hirt M-7 Ornis
The Schempp-Hirt M-7 Ornis aircraft is a single-engine two-seat
(side-by-side) braced high-wing plane of the classical conception.
Wing of rectangular platform is made of wood and fabric. Fuselage
has its frame structure of welded Cr-Mo steel tubes.
In the table you can see natural frequencies of plane wings for weight
450 Kg. The system got good data up to 65 Hz. Higher frequencies have
significant noise. In pictures you can see deflexion line of the wings in
25% points.
During these tests was took that most of antisymmetric modes and
symmetric modes higher that the third were above measuring range.
Vice-versa torsion modes, up to the second, were presented always, frequently in combination with 2nd bending. It was detected, the front-toback modes were important too.
42
C Z E C H A E R O S PA C E P R O C E E D I N G S
1st. symmetric bending mode
relative deflexiion
Tab. 1 — Wing symmetric modes
Faeta
M-7 Ornis
VL-3
Allegro
Tab. 2 — Wing antisymmetric modes
-1
Objectives of further work
-0,4
-0,2
0
0,2
0,4
0,6
0,8
1
Fig. 10 1st
— antisymmetric
1st antisymmetric
bending mode
bending mode
1
Faeta
M-7 Ornis
VL-3
Allegro
wing span
reletive deflexion
relative deflexion
-0,6
0,5
Fig. 8 — 1st symmetric bending mode
bending
mode
Fig. 9 —2nd.
2ndsymmetric
symmetric
bending
mode
-0,8
0
wing span
The next steps are to get higher measuring accuracy in connection
with the structure exciting and efficient techniques for export to the
aeroelastic software. Reducing measuring time is unlikely. It is determined by the capacity of the system.
-1
-0,5
Faeta
M-7 Ornis
VL-3
Allegro
-1
-0,5
0
0,5
1
wing span
Optimization of Stiffened Panel — Design of
Testing Equipment
Optimalizace vyztuženého potahu — návrh zkušebního zařízení
Ing Miroslav Pešák,
Ing.
Pešák Prof.
Prof Ing.
Ing Antonín Píštěk,
Píštěk CSc.
CSc / Institute of Aerospace
Engineering, Brno University of Technology
The questions
q
of the bucklingg of thin walls structures (stiffened
(
panels)
p
) are solved for a longg time byy manyy well-known
scientists. A lot of theories and design processes have been found to predict a buckling and stability of stiffened panels.
Nevertheless even today we are not able to ensure absolutely precise buckling prediction of the stiffened panel. It means
that it is still very important in design stage to verify and test the theoretically calculated data.
The experimental results are very valuable not only to verify the theoretical calculations but also to find and specify
coefficients and last but not least this data are useful to creation a graphs and relations. Very similar reason to
experimental verification of compression of stiffened panel is also in our case. The experimental data will be used for
debugging for the optimization software OPTPAN [1].
Introduction
The design and calculation of the stiffened panels is closely associated with their experimental testing. Probably the biggest experimental
testing of the stiffened panels was done after second world war.
Extensive research in the area of buckling and stability of thin wall
structures made both world aircraft leaders — Soviet Union with the
most famous names: Timoshenko, Hertel, Lipin, Goodier. In USA the
most tests and research work was done at NACA. The main names of
authors of this date: Gerard, Becker, Kroll.
All this measurement gives very valuable data about buckling and
stability of metal thin wall structures.
43
L E T E C K Ý Z P R AV O D A J
Experimental verification of theoretically determined results is still
very important design stage. In spite of the progress of computer
simulating processes in nowadays. This is also the goal of this paper.
This paper continues in extensive research of optimally stiffened
panels which are forced by compressive and shear loading. The first
paper of this research was introduced at CLKV 2006 with the title
Optimalization of stiffened panel with the help of mathematical programming [1] this paper presented a system of computer programs to
optimize and analyze skin-stringer panel for fuselage and lifting surfaces (wing and empennage structures) which was developed at the
Institute of Aerospace Engineering, FME, BUT. This article deals with
the design of the testing equipment for compressive loading of stiffened panel. The experimental results play a very important role in
debugging of optimizing program.
Description of the testing
The goal of the testing is to find a critical panel loading and critical
stresses of the main panel's part. It means that we are trying to exactly estimate the time when skin and stringers buckles and in which
form(local skin or stringer buckling, global buckling of panel, combination of both wrinkling, twisting…). The testing was made in laboratory of Institute of Aerospace Engineering, BUT. Testing was aimed
for compression loading of stiffened (riveted) skin-stringer panel. The
main goal of this testing was verify of functionality of testing equipment.
The main components necessary for experimental verification:
● Stiffened panel
● Testing machine (compressive loading machine)
● Testing box
● Measuring instruments
3/2007
And therefore they are usually incapable of carrying much lateral
load.
The stringers and rings or ribs are attached to the skins by lines of
the rivets, spot welds or perhaps bonding. These joints will be called
upon to transmit forces mainly along their length. Forces parallel to
the skin and directed at right angles to the stringers or rings or ribs
will be limited by the torsion flexibility of these members. Forces normal to the skin will be limited in magnitude by the small bending
strength of the skin and stringers. The primary function of these joints
is thus the transmission, by shear forces, of direct loads in the reinforcing members to the skin and vice versa. Their secondary functions
are indeed essential to the working of the structure but do not give rise
to such large loads.
Stiffened panels are mostly used in the aircraft structure with higher loading factor. Typical utilization of stiffened panels we can find
for example in structure of the L-410 and L-610 airplanes. For optimization as well as for testing there were chosen the dimensions of the
stiffened panel common for such aircraft. The dimensions were also
determined with regard to the maximum testing machine available
space and the maximal compressive force.
External dimensions of the stiffened panel are specified:
A = 800mm
B = 600mm
Loading direction and panels dimensions is shown in Fig. 2
Fig. 1 — Stiffened paneel
Fig. 2 — E
External
t
l panel's
l' di
dimensions
i
and
d the
th di
direction
ti off th
the lloading
In the Fig. 3 you can see the Catia model stiffened panel vs. real
panel.
Stiffened panel (see Fig. 1) is composed of following parts:
Longitudinal reinforcing member
This is represented by stringers and longerons of fuselage shells and
the spar flanges of wings. They are able to carry appreciable tensile
loads and, when supported, compressive loads as well. They can
carry small secondary bending loads, but their bending rigidity is
negligible. So it is customary to describe them as direct load carrying
members.
Skin
Like all thin shells, this is best suited to carrying load in its own surface as membrane stresses. Tensile, compressive and shear loads can
be carried, but reinforcement (lateral support) is required for all but
the first. The thin skins used in aircraft can only sustain and transmit
normal pressure over very short distances by bending. Pressurization
loads in a circular section fuselage can, however, be taken by hoop
tension stresses.
Transverse reinforcing members
These are the rings, frame, bulkheads or diaphragms of fuselages and
the ribs of wings. In the design of these members most attention is
paid to providing stiffness and strength in the plane of the member.
Fig. 3a — Catia model stiffened panel
Testing machine
Basic technical parameters of testing machine:
● Maximal acting force 500 000N
● Production year
● Weight
44
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 6a — The
real testing box
Fig. 6b — Catia
model testing
box
Fig. 3b — The real stiffened panel
Fig. 4 shows the view on testing machine.
All the tests are made on this testing machine (see Fig. 4). This
machine is composed of heavy stand and the hydraulic gear including
dynamometer. These components are connected to the central computer which recording the data from the measurement. Closely after
measurement we are able to gain a relation between translation and
acting force during compression (see Fig. 5).
In the flanges of the ribs and flanges of the spar there are attached
riveting nuts for easier change
another panel. It is also possible
to fasten two stiffened panels on one box (both sides are ready for
panel´s clamping) at one moment by threads. It allows us the faster
change of the stiffened panel to next measurement. Dimensions of the
box have to correspond to dimension A, B of stiffened panel.
Measuring instrument
There are several possibilities how to measure (find) buckling of the
panel's components. We are now in the early stages and we tried to
use only some of them.
1) View ”method“
It is not method to all intents and purposes. Despite of this the
method was used at first test of functionality of testing equipment. This testing was done without any special measuring equipment. Only dynamometer which is situated on hydraulic gear
was switched on.
It is possible to see the waves on the skin caused by buckling
and also local buckling of the stinger without special measuring
equipment.
This method is suitable only for preliminary design and to verify functionality of system because it is very inaccurate. Some
picture from this testing you can see on Fig. 7.
panels
Stabilita vyztu eného panelu
180
160
140
120
Síla[kN]
100
80
60
40
Fig. 7 — Skin vs. sttringers buckling
20
0
-1
0
1
2
3
4
5
6
deformace [mm]
Fig. 5 — Relation between translation and acting force
The curvature changes of the curve on Fig. 5 shows the buckling of
individual components of stiffened panels. The first curvature change
about at 20kN shows the skins buckles and only stringers and box
flanges carry over the increasing loading. Next curvature change
about 120kN shows lost the stringer's stability. It is the estimated loading when panel buckles.
Testing box
The testing box belongs to the fundamental part of the testing equipment. The Catia model vs. the real box is shown in the Figure 6. This
special box enables us to provide a realistic simulation of the boundary conditions during the testing of the stability of the stiffened panel.
2)
Buckling measurement by digital camera
The basic of this method is to harmonize camera and loading
time. The principle of this method is to picture the stiffened
panel by video camera and next evaluate a translation of the
selected point (like movement of a pixels). For better orientation the panel was divided by grid see Fig. 8.
For better view the camera was situated on the bottom of the panel
a picture was done by the system of the mirrors. See picture 9. where
you can see obvious differences of the unloaded panel vs. loaded
panel.
This method is not very useful for solution problem of starting
panel's buckling. The lines that makes a grid were very thick and does
not correspond with pixel's magnitude. So recording exact time when
45
3/2007
L E T E C K Ý Z P R AV O D A J
Fig. 9 — Unbuckled vs.. buckkled stiffened panel
Fig. 8 — Grid on the stiffened panel
skin starts buckle is quite difficult. Also evaluation of this method is
very time consuming. This method is more suitable for measuring
quicker experiments with bigger shape deviation.
3)
Testing by drift meter
There were used a drift meters with designation Sylvac to find
the critical compression loading of stiffened panel components.
The drift meters are very sensitive equipment which is able to
record even a very small movement of the skin. From the theory of a buckling of thin plates it is clear that skins buckles in
a waves. In this theory is necessary to find a best drift meter
position. It means to situate the drift meter to the top of sinus
curve (not to the nod!). The establishing the correct position is
done with the help of Fig. 10.
The positioning of the drift meters on the skin of stiffened panel
is shown on Fig. 11 Buckled vs. unbuckled skin.
F 10 — Number of
Fig.
waves on the skin at
w
ccompressive loading
Nowadays it was used only thre drift meters to verify functionality of
this method. Despite of it we reached very valuable result data see.
Fig. 12. This method of measuring seems very attractive and we will
make next measurement with more drift meters which allow us to
measure all important areas in skin field.
Conclusions
The design of the testing equipment and possible measuring techniques for evaluation of the testing a stability of the stiffened panel are
introduced in this paper.
The ultimate dimensions of the stiffened panel were determined
with regard to the possibility (maximal allowed force of testing
machine is 500kN) of the laboratory of Institute of aerospace engineering. For this designed panels was made a ”testing box“ which is able
to fix two stiffened panel at the one moment. The simulation of the
real boundary condition during the compression loading is the main
goal of this special testing box.
There were presented three methods to evaluate the results at stiffened panel testing in this paper. The last one — the testing by drift
meters looks to be the most efficient. This method will be develop
also in the future and we hope that when will be used more drift
meters in one skin field and we will be able to perfectly define the skin
behavior during compression loading.
The next our effort will be to find and verify another measuring
method eg. by tensiometers, aripots. …
As soon as the most effective the measuring method will be found
and we will be sure that the experimental results are accurate we can
start the next stage of the extensive stiffened panel research. In this
stage it is planed testing of large number of stiffened panels optimally designed by the help of OPTPAN [1]. The measured data will be
retroactively used to debugging of the OPTPAN.
Fig. 11a — Starting measurement Fig. 11b — Buckled skin
nonbuckled skin
1
0,5
0
0
10
20
30
40
50
60
70
80
90
100
-0,5
-1
Øada1
Øada2
Øada3
-1,5
-2
-2,5
-3
-3,5
Fig.
g 12 — Relation between skin deviation vs. loading
g force
References:
Pešák, M., Píštěk, A.: Optimalization of Stiffened Panel with the
Help of Mathematical Programming, Praha CLKV2006, 2006
[2]
Píštěk, A.: Kandidátská disertační práce, Brno, 1980
[3]
Michael C. Y. Niu: Airframe Structural Design - second edition,
1999
[ 4 ] Píštěk, A., Hobza, P.: Optimum Design of Stiffened Panel Using the
Method of
[ 5 ] Kopřiva, Z., Hotovec, Z., Morkus, V.: Experimentální ověření optimalizace vyztužených panelů konstrukce křídla, Brno, 1987
[ 6 ] Gerard G.: Handbook of Structural Stability, Part 5 — Compressive
Strength of Flat Stiffened Panels; National Advisory Committee for
Aeronautics — TN 3785, 1957
[1]
46
C Z E C H A E R O S PA C E P R O C E E D I N G S
Design and Manufacture of Aircraft Parts
by LF-Technology
Konstrukce a výroba součástí letadel technologií LF
I
Ing.
(Chem.
(Ch
E.)
E ) Lukáš
L káš Křípal
Kří l / Institute
I tit t off Aerospace
A
Engineering,
E i
i
Brno
B
University of Technology
The present paper discusses
i
a LF-Technology ((Letoxit
i Foil)
i ) used ffor manufacturing
f
i off 33D aircraft
i
f propeller nose cone
part and leading edge nose rib. LF-Technology was introduced by 5M Ltd. from Czech Republic. LF-Technology is
modified RFI technology (Resin Film Infusion). The simple and fast process is based on vacuum and/or press-assisted
resin film impregnation of a dry reinforcement and core material in the mould, with high-temperature curing
(autoclave isn't necessary). The process allows placing required type and number of layers to different locations, and
to achieve the exact resin content required as a function of the number of resin layers used. Only one type of resin
film is necessary, and this can be combined with almost any type of reinforcement. The influences of a number of
processing parameters on the quality of cured parts are discussed along with previous flat sample testing results. LF
Technology gives freedom to designers, helps push down prices of composite products and increases properties and
reliability, which is supported by favourable responses from customers and quick and relatively easy development of
new products. All these advantages promise very good future for LF Technology.
Tento článek pojednává o LF Technologii (Letoxit Foil) aplikovanou na výrobu krycí části propeleru (kužel a náběhové
žebro. LF-Technologie byla vyvinuta firmou 5M s.r.o. Česká Republika. LF-Technologie je modifikovaná RFI
technologie (Resin Film Infusion). Jednoduchý a rychlý proces je založen na aplikaci vakua (volitelně i vnějšího tlaku)
na kladenou suchou výztuž kombinovanou s epoxidovou pryskyřicí ve formě folie (použitelnost technologie i pro
sendviče). Takto uspořádaný systém je poté vytvrzen zvoleným teplotním cyklem bez nutnosti použít autokláv. Proces
umožňuje zvolit požadovaný typ a počet vrstev a dosáhnout tak přesného množství pryskyřice (”šití materiálu na
míru“). Je možno použít pouze jeden typ pryskyřice, který se dá kombinovat s téměř každým typem výztuží. Vlivy
procesních parametrů na kvalitu vytvrzených kompozitů jsou zde diskutovány spolu s předešlými výsledky testování
rovinných jednosměrných vzorků. LF Technologie dává svobodu designérům, snižuje výrobní náklady a zvyšuje
vlastnosti a spolehlivost takto vyrobených kompozitů, což je ostatně podloženo ohlasy zákazníků a díky rychlému
a relativně jednoduchému způsobu jak vyrobit nové produkty. Všechny tyto výhody slibují dobrou budoucnost pro LF
technologii.
Keywords: LF-Technology, mechanical testing, RFI technology.
Introduction
Infusion manufacturing techniques are gaining popularity in aerospace applications. It is about 20 years from first introduction of solid
epoxy resin film and its technology named resin film infusion (RFI).
Czech producer and provider 5M Ltd. makes high-quality specially modified epoxy resin films and offers own modification of RFI
technology called LF Technology, named after epoxy resin marking
Letoxit Foil. This technology was described in detail on previous
Reinforced Plastics International Conference 2005 and CLKV 2006
meeting. Our target was the verification of LF technology potentials
on the most complex shape.
A spinner nose dome made by LF technology
5M Ltd., a Czech company active in the field of composites and
sandwich materials, introduced 2004 the LF Technology (Letoxit
Foil) for dry composite manufacturing. The one-shot process is based
on a vacuum and/or press-assisted resin film impregnation of a dry
reinforcement and core material in the mould, with high-temperature
curing. LF Technology is based on laying dry reinforcement and core
material to the mould with layers of foil polymer material, known
under commercial name Letoxit Foil. Whole composition is vacuum
bagged and cured at elevated temperature. The scheme can be seen in
the Fig. 1.
The most common curing method (heat) is used. The consolidated
laminate is placed in oven and the heat provides energy which drives
the cross-linking of the matrix to completion.
The Institute of Aerospace Engineering owns facility for composites production by LF-Technology is shown on Fig. 2. It consists of
Fig. 1 — Scheme of RFI (LF) Technology
Fig 2 — LF technology facility (below)
Fig.
47
L E T E C K Ý Z P R AV O D A J
a vacuum pressure source; ejector vacuum pump and an intelligent
programmable hot air dryer.
3/2007
Curing cycle
Used Materials
a) For a mould processing
- E-glass - twill weave, 160 g/m2 a 280 g/m2
- Epoxy gelcoat P for moulds /R+G/
- Gelcoat hardener VE 3261
- Epoxy resin Martens-Plus-EP / R+G/
b) For the spinner cone
- E-glass - twill weave, 160 g/m2 a 280 g/m2
- Epoxy resin (film) LFX023
- Basic wax release agent
- PVA mould release agent
- Epoxy gelcoat clear UV stable Novalis T 100 A
- Hardener Novalis T 100 B for gelcoat Novalis T 100 A
At first 3D CATIA V5 model was created and coated with twill texture. Interferences appeared and this was very helpful in lay-up
design (see Fig. 3). We had cut four uniform layers to fully cover all
inner mould area. At first time only three parts were applied with no
success because of resin layer was pre-melted on fabric surface (partial loss of twill style ability).
Fig. 3 — Twill fabric interferences on cone model (side view) and
spinner cone on ultra-light
ultra light aircraft
Fig. 6 — Selected curing cycle
Mould construction
It was necessary to make a high temperature resistant one piece
mould. The female part was made from E-glass twill fabric and Martens epoxy resin system.
For a mould making a male part was used (see Fig. 7). The mould
surface was cleaned and after rubbed with PVA mould release agent.
Then epoxy gelcoat P (Fig. 8) for moulds was applied and layers of
E- glass reinforcement were stacked on mould surface. Firstly light
weight fabric and then 3 layers of the heavier one (see Figs. 9 and 10).
The twill fabric was trimmed angle-wise 45° for better handling.
Whole composition was then wet-out by high performance
MARTENS PLUS EP epoxy resin system (see Fig. 11) and cured, see
Fig. 12.
Processing parameters
Important factors during processing are: high vacuum pressure level,
fresh resin film and sufficient application of mould release agent
(120° C resistant).
For the spinner curing cycle with slow heat up rate and dwell has
been used. This curing cycle is shown on Fig. 6 signed as J. The
curing temperature 120[° C] was applied with 60 minutes of curing
time altogether with vacuum pressure about 80[kPa].
Fig. 7 — Spinner cone
master model
Fig. 8 — Epoxy gelcoat P
application
Fig. 9 — The first low weight
fabric layer
Fig. 10 — The second heavy weight
fabric layer
Fig. 4 — Typical RFI sample lay-out
Lay-up & Infusion flux
Here, ply-by-ply lay-up was chosen, see Fig. 4. Infusion flux scheme
is shown on the picture below. Here, resin must travel at most a few
millimeters to fully wet out the reinforcement.
Fig. 11 — Stacked & wet-out reinforcement Fig. 12 — The mould curing
48
C Z E C H A E R O S PA C E P R O C E E D I N G S
MARTENS PLUS EP
Solvent- and filler-free 3-component epoxy laminating resin for hightemperature components (approx. 230° C) with high static and
dynamic strength.
Mixing ratio:
100 : 100 : 1.5 parts by weight of resin to hardener to accelerator.
Processing time: 12 hours at 25° C or 2 hours at 50° C.
Curing until demoulding 24 hours at 80° C.
Annealing 24 hours at 100° C + 10-15 hours at 230° C.
Demoulding is possible at the earliest after approx. five hours of
annealing at > 80 °C.
Spinner cone processing
The side and the top view on the mould with displayed dry reinforcement and resin film layers can be seen on Fig. 13. The following figures (14 and 15) present application of vacuum bag over the mould and
final stage of spinner cone [3].
Before reinforcement and resin film are laid into the mould is good
to apply slim gelcoat layer to get smooth pinhole free surface. In this
case is really necessary to choose suitable gelcoat system matching
the epoxy resin film system.
flanges round its edges so that it can be glued to the wing skin and
spar webs.
The technology has already been successfully applied to a composite spinner nose cone component.
A wooden model for a male mould has been chosen. Then a PVA
mould release agent was applied on release tape (Fig. 16).
For the first attempts were E-glass twill fabric 200 g/m2 and epoxy
resin LFX023 200 g/m2 used. E-glass cloth was chosen on purpose,
to observe saturation effect.
Final part was completely saturated but on the upper surface asperity appeared due to the excess of the resin.
On the final carbon fibre rib resin rich areas they were ground off
and whole upper surface with clear gelcoat layer has been covered.
The resin excess rich areas appeared mainly there, where additional
layers of tear-off fabric, perforated foil and bleeder were placed. It is
possible to cut and place these layers on flat parts and on whole mode
put tear-off fabric.
In this attempt a 285 g/m2 E-glass fabric has been used. However,
locally unsaturated areas appeared. It was due to the shorter curing
cycle where resin during flow did not have enough time to impregnate all fibres in composition. For the next step modified curing cycle
has been applied.
Fig. 17 — Resin placement
Fig. 13 — The side and the top view on mould with dry reinforcement
and resin film layers
Fig. 18 — Composition curing under
vacuum
Fig. 19 — Upper side view on
final part
Final carbon model (Fig. 19) was cured (Fig. 18) under the longer
cure cycle with ramp (2° C/min.) and twenty minutes dwell.
Fig. 14 — Vacuum bagging
Fig. 15 — Finished spinner cone
Leading edge nose rib
It was necessary to make a structural part to prove the possibilities of
LF technology. For this purpose a nose rib section has been chosen.
The rib has been formed in four parts from carbon fabric by vacuum resin infusion technology using solid sheet foil resin. This rib has
Fig. 16 — Male wooden mould
LF Technology applications
The LF technology was successfully applied on VUT 100 COBRA —
sandwich baggage hold wall; M 101 Expedition airplane — interior
parts (door panel, spring bracket, right and left refreshment cabinet
etc.); engine hood of ultra light airplane CHS 701SP (Zenair), Carbon
knee orthotics (ING Corporation), bicycle frame parts etc. [4]
Conclusions & Summary
The main field of RFI technology is in aerospace technology where
this method can pull down cost of finished composite.
The LF technology has shown really effective way how to process
high performance composites. Because of the shape difficulty and
verification of LF technology the spinner cone was chosen. Just for
reference only, preparation of the mould takes here 40-50% less time,
so it is 3 times faster with LF Technology than with the hand lay-up.
In contrast to the classic hand lay-up technique, here we obtained
lighter part with less resin content and smoother surface.
49
If we have perfectly prepared mould surface we can apply gelcoat
as a first surfacing layer or directly apply resin film layers. Here is one
problem, without a gelcoat, pinholes on the surface can appear.For
the perfect saturation of whole composition an essential role plays
a heat-up rate (ramp), dwell in the cure cycle and also resins freshness. Before use, epoxy resin film coil has to be defrosted at 25° C for
5 or 6 hrs. (12 hrs. is better) to remove residual humidity. The resin
single layers can be pre-melted by heat gun for better handling.
Use of LF technology is quick and with its help we can ”tailor“
final part, nevertheless it is necessary to be an experienced and skilled
worker to be able to manufacture more complicated shape parts.
Considering this, for every new piece it is necessary firstly to try
out error-free manufacturing process and input contingent improvements.
To make a rib model for frequent use it is favourable to select
a different mould material e.g., metal or modelling board material
RenShape© BM 5055 (heat resistant up to 140° C) or SikaBlock®.
It wasn’t possible to make a female mould tool from wooden male
one, so that’s why these results are not described here in detail but
probably the same imperfections, like with male mould tool, would be
expected.
L E T E C K Ý Z P R AV O D A J
3/2007
Two piece mould tool to be made from materials mentioned above
is recommended here. That’s the way we can obtain both sides of final
product perfect and good compaction of dry reinforcement stack and
epoxy resin film layers can be achieved.
The results of this work has been published and presented on national and international conferences. The results of nose rib made by
LF technology are presented for the first time ever.
References:
[1]
C. Williams, J. Summerscales, S. Grove: Composites Journal, Part
A 27 (1996) 517
[2]
Pavlica R., Eder M.: RFI - Key for optimised Composite Structures, Reinforced Plastics 2005 (XXIII International Conference),
24th to 26th May 2005, Karlovy Vary, Czech Republic - conference proceedings, ISSN 1214-6412
[3]
Křípal L.: Composite Propeller Spinner Nose Cone Made by LFTechnology, CLKV Scientific Workshop 2006, Prague, 2nd - 3rd
November 2006 - oral presentation
[4]
Křípal L., Polymer Fiber Reinforced Composites Processed by RFI
(Resin Film Infusion) Technology - Comparisons of Mechanical
Properties and Applications; 13th Conference of Sociedade Portuguesa de Materiais - MATERIAIS 2007, 1st - 4th April 2007, Porto,
Portugal
Ducted Fan Power Unit Demonstrator
for Ultra Light Airplanes
Demonstrátor ventilátorového pohonu pro ultralehká letadla
Ing.
g Pavel Růžička / Department
p
of Aerospace,
p
, Czech Technical Universityy
of Prague
The paper deals with the design of the ducted fan power unit demonstrator for ultra light airplanes. The demonstrator
is intended for experimental verification of power unit design concept and data acquisition, which will be used for optimization of structural assemblies, dynamic tuning and gaining power unit operating characteristics.
The paper is focused on the base frame and the engine frame design, thrust measuring of the power unit, engine mounting framework design, as well as on modifications of the rotor system for real stress measuring of the transmission
shaft and measurement of the rotor system vibrations. Furthermore, on measurements of thermal emissions, possibility
to use this exhaust energy for the engine bay ventilation and increasing thrust in the jet nozzle.
Příspěvek se zabývá konstrukčním řešením demonstrátoru ventilátorového pohonu pro ultralehká letadla určeného
pro experimentální ověřování konstrukčních řešení pohonu a získávání dat pro následnou optimalizaci konstrukčních
uzlů, ladění systému jako celku a získání funkčních charakteristik pohonu.
Příspěvek je zaměřen zejména na popis a funkci základního a motorového rámu při měření tahu pohonné jednotky,
motorového lože s měnitelnou tuhostí a variant rotorového systému pro měření reálného zatížení transmise a vibrací.
Dále pak na měření emisí tepla a využití energie spalin ve výstupní trysce a pro ventilaci motorového prostoru.
Keywords: UL aeroplane, cold jet, ducted fan, ultra-light rotor design, engine mounting, engine
test stand.
Introduction
The ducted fan power unit demonstrator (Fig. 1) is an engine test
stand for correct function verification of the new power unit with the
ducted fan powered by a piston engine developed for ultra lightweight
airplanes. The power unit design and installation required the relatively long transmission shaft driven by the superbike piston engine
Yamaha YZF R1. The engine was chosen for its favourable powerweight ratio. The weight is a limiting factor in this ultra lightweight
aircraft category, therefore all the aircraft structure is designed for the
minimal weight. The power unit structure is an unusual concept, the
engine was not initially designed for employing in airplanes and all
power unit parts must be optimized to decrease weight. Therefore, it
is necessary to test the power unit in operation and obtain the data needed for the optimization of all the system by measuring step by step
and finally prove its correct function.
Fig. 1 — The
Fi
Th d
demonstrator
t t
50
C Z E C H A E R O S PA C E P R O C E E D I N G S
Problem description
The fundamental requirement on the power unit is the useful thrust
for the aircraft. The thrust measurement of the power unit is a major
task for the demonstrator. The dual frame of the demonstrator is used
for this purpose. The frames are connected by a metal strips equipped
with strain gauges. The thrust can be measured during the power unit
operation after the appropriate calibration.
The superbike engine mount is rather different then the aircraft
engine framework. That is the reason for a special design of the engine mount. The demonstrator beam frame represents a sufficient rigid
structure to simulate rigid mounting of the power unit. In case that the
power unit is installed to the non-rigid aircraft structure or silent
blocks are used it is necessary to simulate non-rigid mounting of the
power unit. The demonstrator engine mounting allows to insert various silent blocks. The mode of the power unit mounting affects dynamic behaviour of the rotor system, which is excited by the piston engine. It is important to choose the proper stiffness of the rotor bearings
based on the vibration measurements. The unbalanced rotor system
can be a source of additional vibrations. These vibrations are transferred to the frame or aircraft structures. It is necessary to know which
vibrations really appear in the structure and under what operating
conditions. For this purpose, the demonstrator is equipped by accelerometers for measuring vibrations under working conditions.
It is difficult to calculate this engine excitation particularly if the
pressure-volume diagram is not available. This diagram can be obtained by a modified plug or it is necessary to modify the engine. The
calculation of forces in a crank gear is still affected by errors because
the mechanical efficiency of automotive gears depends on many
quantities (temperature, rotation speed, oil properties, etc.). The character of the piston engine function produces peaks of engine torque.
The peaks can exceed mean torque measured by the dynamometer by
the order of hundred percent. These peaks do not affect the engine
power but add additional loads to the shafts. If weight limits cannot
be exceeded it is necessary to know exactly the stress of the rotor
parts and design the parts precisely for this stress only. The transmission shaft and transmission shaft accessories represent important
weight in the power unit assembly. The demonstrator is suitable for
this strain gauge measurement of the transmission shaft under operating conditions. The measured data are useful for rotor parts optimization and dimensioning of the elastic couplings that can improve the
dynamic behaviour and misalignment of the rotor parts.
The demonstrator also allows aerodynamic measurements of the
power unit flow path and so it is possible to compare the measured
and the calculated data. The data serve to optimize the exhaust nozzle
including the cooling system.
The heat balance of the power unit is very important. The heat generated by the engine has to be effectively conducted away from the
engine bay and utilized if possible. The temperature can be measured
at all the important places of the demonstrator, such as bearings, contact seals and at the simulated enclosed aircraft engine bay. Moreover,
the course of temperature along all of the exhaust way can be monitored. The knowledge of this temperature allows optimization of the
rotor mast attachment point and the airplane body structure and properly using of the structure materials. Mainly mechanics of composite
materials may be temperature-dependent. The engine bay cooling by
the exhaust duct ejector can be tested with the help of the demonstrator as well as a thrust gain of the exhaust gas in the jet nozzle.
Propulsion unit
The propulsion unit consists of the four-cycle in-line four cylinder
engine Yamaha YZF R1, model 2004. The stroke volume is 998 cm3
and the power is 134 kW at 12 500 RPM. The fan is driven by the
transmission shaft. The airflow is guided and prewhirled by two composite inlet channels and then flows out through the jet nozzle that
allows exhaust heat utilization. The designed power unit was described in [1].
Ducted fan power unit demonstrator
The demonstrator (Fig. 2) consists of the basic frame and the engine
frame with the fan sliding support as shown above. The basic frame is
a rigid part fixed to the ground. It can be easily fixed into the anchor
plate in a test room. The engine frame represents an aircraft body and
serves for the power unit trial installation. The engine mounting is
mounted in the front and the fan sliding support that allows installation of different length transmission shafts is at the rear. Both the frames are joined by four accurate metal strips (Fig. 3). The strips are
installed to achieve minimum stiffness in thrust direction. There is
a strain gauge full bridge installed on each metal strip (Fig. 4). Measured deformations of the strips under the applied thrust are evaluated
with the help of a calibrating curve. This calibrating curve is obtained
by thrust meter measuring (Fig. 5) applying hydraulic cylinder (Fig.
6) force on the demonstrator in thrust direction and comparing it with
the strain gauge data.
Fig. 2 — The basic frame and the engine frame
Fig. 3 — The metal strip
Fig. 4 — The strain gauge full
bridge
Fig. 5 — The thrust meter
Fig. 6 — The hydraulic cylinder
The spherical engine mounting corresponds with the schema ”from 4
to 3 junction points“ as shown in [2]. The engine mount design tries
to respect last trends in this area and provides certain instruction to
ultra lightweight aircraft designer which has some problems with
installation of the motorcycle engine to the airplane, though the engine mounting is advisedly overdesigned for testing purposes. Thus, the
51
L E T E C K Ý Z P R AV O D A J
3/2007
engine mounting can simulate nearly rigid structure (Fig. 7).
Nevertheless, a set of three joints allows one-sided or two-sided inserting of the silent blocks with various dimensions (length and diameter) (Fig. 8). That way it is possible to change the engine mounting
stiffness within the satisfactory range. The optimal silent blocks combination can be easily detected with the help of vibration measurements by a set of installed accelerometers (Fig. 9). It is possible to
decide what stiffness of the engine mount leads to decreasing vibrations and if the silent blocks must be used on the basis of these vibration measurements.
Figure 10 — The transmission shafts
Fig. 7 — The engine mounting
Figure 11 — The technological
transmission shaft
ture of the fan bearings. The fan bearings are sealed and greased for
whole life due to decreasing weight and simplicity. The knowledge of
the bearing temperature allows selecting proper material of the bearing body with regard to weight reduction and function.
Fig. 8 — The silentblock
Fig. 9 — The accelerometers
The relatively long transmission shaft (Fig. 10) is a critical part of the
rotor system. The shaft is designed as a tube with a large diameter and
a very thin shell to get good weight-stiffness ratio as shown in [3].
Another cut of the weight has been achieved using a hybrid transmission shaft instead of the metal one. There is inserted composite tube
(carbon-epoxy) between the metal flanges as shown in [4]. Two flexible couplings protect the transmission shaft from the additional
loads besides the torque and improve dynamic behaviour. The flexible couplings align possible misalignments of the engine rotor parts
and the fan. The transmission shaft and the accessories represent considerable weight and require optimization. The torque peaks can be
best obtained by strain gauge measuring under operating conditions.
That means the design process has to be iterative. It stands to reason
that dynamic problems rise with increasing of the length and weight
of the rotor system (transmission shaft). Therefore, the short modified technological transmission shaft (Fig. 11) is used for this measuring. The flexible couplings are substituted for duralumin adapters of
the same length. The hybrid shaft is not used because the strain gauges can be better applied on the smooth metal surface. The fan sliding
support is equipped with an adjusting mechanism for precise axis
alignment so the flexible couplings can be omitted. The strain gauge
data transmission under rotation must be done wireless.
Another important data can be obtained during these tests, for
example correct function of the seals in the new gearbox housing
cover (Fig. 12) and pair of engine output shafts or running tempera-
Fig. 12 — The gearbox housing cover
The thermal emissions must be known for the power unit installation
into the closed airplane body and the cooling system including the
engine bay ventilation must be correctly designed. It is possible to
determine energy distribution from fuel to mechanical work, radiator
off-heat, exhaust heat and heat radiated into the airplane body by temperature measuring. Based on the measurement it can be evaluated if
the exhaust gas energy is usable for the ejector engine bay ventilation
(Fig. 13). The metal jet nozzle (Fig. 14) of the power unit allows to
guide exhaust gas to the air flow behind the fan. The possible thrust
gain can be directly measured by the strain gauges applied on the
metal strips and then evaluated in terms of weight reduction. The
mounting hole in the jet nozzle body can be used for aerodynamic
measurement. It is possible to utilize the radiator off-heat by the same
way as shown in [5].
52
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 14 — Themetal jet nozzle
References:
Fig. 13 — The ejector engine bay ventilation
[1]
Růžička, P., Hanus, D.: Feasibility study of Ultra-Light Cold Jet
Aircraft Propulsion Unit and Rotor System Design; In Letecký
zpravodaj, November 2004, No. 3/2004, pp. 16-19. ISSN 1211877X
[2]
Benda, L.: O kmitání pohonných jednotek 1, 2 and 3; In Pilot LAA
ČR, April-June 2006, No. 4/2006, pp. 18-19; No. 5/2006, pp. 1213; No. 6/2006, pp. 12-13
[3]
Růžička, P., Poul, R., Hanus, D.: Propulsion Unit for The Ultralightweight Airplanes Design and Optimization of the Transmission
Shaft; In 5th International Conference on Advanced Engineering
Design, 11-14 June 2006, Prague. AED, 2006. AED CD-ROM.
ISBN 80-86059-44-8
[4]
Poul, R., Růžička, P., Hanus, D.: Design of Carbon Composite
Driveshaft for Ultra-light Aircraft Propulsion System; In Acta
Polytechnica, Vol. 46. May 2007, No. 5/2006, pp. 40-44. ISSN
1210- 2709
[5]
Strnad, O.: Zadní část malého sportovního letounu; Praha, 2002.
72 s. Diplomová práce na FS ČVUT na Odboru letadel. Vedoucí
diplomové práce Robert Theiner
Conclusions
The ducted fan power unit powered by the piston engine is an unusual design of an aircraft propulsion unit. Therefore, it is necessary to
obtain the field data for the structure optimization and verification of
design characteristics. In this respect, the demonstrator is essential for
the power unit development.
Acknowledgements
This research was carried out at the Aerospace Research Centre supported by the Czech Ministry of Education, Sports and Youth, by the
Grant No. LN 00B 051.
The Methodology of Winglet Aerodynamic
Design for UL and VLA Aircraft
Metodika aerodynamického návrhu wingletu pro UL a VLA letoun
Ing Robert Popela Ph.D,
Ing.
Ph D Ing.
Ing Pavel Zikmund,
Zikmund Ing.
Ing František Vaněk Ph.D,
Ph D
Ing. Martin Kouřil Ph.D / Institute of Aerospace Engineering, Brno
University of Technology
The paper describes
i
ultra-light
i
aircraft
i
f suited
i
winglet
i
optimization
i i i process. Based on C
CFD analysis
i off aerodynamic
i
effectiveness, the geometrical shape of winglet was optimized. Then final design was also analyzed from stability and
loads point of view. Finally, the substitute wing method was used for spanwise lift distribution analytical solution and
was compared to CFD results.
Článek popisuje postup optimalizace wingletu, jakožto zařízení pro zvýšení efektivnosti nosné plochy, se zaměřením na
použití v kategorii ultralehkých a velmi lehkých letounů. Cílem práce byla optimalizace geometrického tvaru s využitím
parametrického CAD modelu a CFD analýzy vlastností návrhu, vyhodnocení efektivnosti a rovněž podrobný rozbor
vlivu wingletu na zatížení křídla. V rámci tohoto rozboru byla rovněž úspěšně ověřena metodika použití náhradního
křídla s vyšší štíhlostí pro výpočty zatížení křídla s wingletem pomocí Glauertova řešení nosné čáry.
Keywords: winglet design, CFD , load distribution, lift distribution.
1. Introduction
The research project target was UL aircraft wing modification for increased aerodynamic efficiency by adaptation of winglet. The first part
of the project was focused on aerodynamic optimization of wingwinglet system. The optimization criterion was the minimal drag.
Then the final winglet design was also analyzed at sideslip flight to
simulate un-symmetrical stall on winglets and to determine its influence on stability and aerodynamic loads.
The second part was focused on spanwise lift distribution determination by analytical approach. It is one of the aerostatic data basic components and is necessary for stress analysis. The substitute wing method was used and verified for calculation of spanwise lift distribution.
Finally, the analytical and CFD method results were compared.
53
L E T E C K Ý Z P R AV O D A J
3/2007
2. Winglets design and analysis
For aerodynamic optimization the winglet fully parametric CAD
model was designed. The following set of parameters was selected for
optimization : position, size, sweep angle, dihedral angle, toe-in angle
and twist. A low-Reynolds Eppler airfoil was selected as a suitable for
low-Reynolds number flight conditions (Re = 2*105 - 1,2*106).
During the optimization process, there were designed couple variants
of winglets and also extended wings. The sample designs are shown
in Fig. 1.
Fig. 3 — Drag change caused by winglets
The sideslip flight analysis of chosen design showed, naturally, that
the left and the right winglet stalls at different flight regime. The
influence of sideslip on lateral-directional stability can be seen mainly at the yaw moment values. The stall came firstly on the winglet at
windward side. It causes drag increase on this side of the wing. The
yaw moment behaviour is shown on the Fig. 4.
Fig. 1 — Sample view of design versions
The unstructured meshes were prepared in the ANSYS ICEM CFD
5.1 software. Domain was bounded by hemisphere, and the final
winglet version domain was bounded by sphere to simulate slideslip
fligt. The computational mesh size was approximately elements. The
CFD software Fluent 6.3.26 was used for computations. All symmetry cases were solved for angles of attack from -2° to 10°. And the
final winglet version was solved in the range of angle of attack from
0° to 14° and angle of sideslip from 0° to 15°.
3. Proces of solution and modification of winglets
At the beginning the first shot winglet was designed and analysed.
The results were evaluated qualitatively and quantitatively. Forces
acting on the wing and winglet were extracted separately. Critical
parameter was forward force component on winglet. After the evaluation, the winglet was modified, and then analysed again, couple
times. The main influence on overall aerodynamic efficiency had
winglet toe-in angle.
4. The final winglet version
The two winglet versions (finally named No. 1. and No. 2.) gave the
most satisfactory results. The comparison of glide ratio of these two
Fig. 2 — Glide ratio for two best designs
versions, original and extended span wing is shown at Fig. 2. Winglet
No. 2 gives the best result for low lift coefficient, but loses efficiency
at higher cL. This is caused by stalling of the flow on the winglet surface at higher angles of attack. On the other hand winglet No.1 works
in wider area, but has the breaking point (wing lift coefficient, where
the winglet drag becomes negative) at higher lift coefficient (see Fig.
3). The winglet with wider efficiency area was chosen as a final
design.
Fig. 4 — Wing
Fi
Wi yaw moment
Results show the efficiency breaking point at cL = 0.35. We can expect
increase of drag at higher speed. But this increase is insignificant against the drag reduction at higher lift coefficient. The aerodynamic
moments character, especially the yaw moment, are not linear because of difference in the stall on winglets. The stall at only one winglet
(windward side) improves lateral-directional stability.
5. The spanwise lift distribution determination by
”substitute wing“ method
The principle of method (see ref. [3], [7]) is substitution of the wing
with winglet by so-called substitute wing with the higher aspect ratio.
The wing substitution is based on extension of the original wing. The
goal is to achieve the same shear force and similar bending moment
at the wing root. This substitute wing is equivalent to the wing with
winglet in term of structural loads, not in terms of aerodynamic efficienty. The spanwise lift distribution of the substitute wing was determined by the Glauert method (software Glauert III, [2]). The acquired spanwise lift distribution was adapted to original wing span (with
winglet) and wing area. In the Fig. 5, there is shown the comparison
between additional spanwise lift distribution of the wing without
winglet, the substitute wing and the wing with winglet. The described
method was compared to CFD analysis results.
6. The aerodynamic loads
The lift distributions over the wing with winglet were determined
from the CFD results. These distributions were compared with the
analytic lift distributions from the lifting line theory for the wings without and with winglets and for the substitute wing. The importance of
the root bending moment and shear force was emphasized.
The surface pressure distribution from CFD results was for each
case exported to standard CFD postprocessor Tecplot format. Then
54
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 5 — S
Fi
Spanwise
i lift di
distribution
t ib ti
Fig. 6 — Wingtip control
slices distribution
Fig. 9 — Wi
Fi
Winglet
l t lift di
distribution
t ib ti
the x axis were calculated by the integration of the unit load in the
z axis direction along the winglet height. The centre of pressure for
winglet was determined and the additive bending moment is given by
the following equation: Mbendingadd(z) = Tz⋅Δy
Δ (z), where: Δy
Δ (z) - vertical distance of the resultant Tz (Fig. 8). Real distributions of ”lift“
along the winglet height are shown in Fig. 9.
7. Results of a load comparison
The Table 1 shows the relative root bending moment for the wing without and with winglets and the relative root bending moment calculated from CFD for the wing with winglets. Data are compared for an
identical shear force in the wing root, therefore for the identical lift
coefficient, because wings had the identical reference area. The wing
bending moment from the analytical lift distribution of the substitute
wing was set as a reference for this comparison (100%).
Fig. 7 — Sample slice pressure distribution
the boundary conditions, wing and winglet, were divided in approximately 62 control slices. At these slices local lift and drag were integrated using in-house developed code. See Figs. 6 and 7 for control
slices distribution and example of visualized pressure distribution in
control slice.
The wing bending moment was determined from the integration of
spanwise lift distribution. The additional bending moment from the
winglet lift was determined in the following manner: The unit shear
force (Tz) in the z axis direction and the unit bending moment about
As it can be seen from Table 1, the difference between wing root bending moment value obtained from CFD analysis and from analytical
Glauert's solution of the substitute wing is less than 1.5% for all cases,
for angles of attack higher than 0deg, it is less than 1%. The used methodology for the analytical lift distribution of the wing with winglets
gives so a very good compliance with load from CFD including the
additive bending moment from the winglet.
Fig. 8 — Additional bending moment generated by winglet
Fig. 10 — R
Fi
Relative
l ti diff
difference off th
the bending
b di momentt
Tab. 1 — Relative wing
g root bending
g moment comparison
p
55
L E T E C K Ý Z P R AV O D A J
3/2007
Also the bending moment values along the wing semispan were
compared and the results are shown in Fig. 10, where the relative difference between local value of the bending moment (based on integration from wing tip) for substitute wing analytical solution and for
CFD analysis can be seen for different angles of attack. The difference is greatest at the tip of the wing, where, naturally, analytical solution have the zero value of the local bending moment. Along the
semispan the difference decreases to values about 1% at the wing
root. See Fig. 11 also for circulation distribution comparison.
8. Conclusions
There was optimized UL aircraft suited winglet from aerodynamic
efficiency point of view. Based on CFD analyses of fully parametric
CAD model the most promising design was chosen and then detailed
analyses of sideslip flight regimes influence on lateral stability and
analysis of aerodynamic loads were accomplished. Based on this analysis, the substitute wing method for fast analytic wingspan lift distri-
Fig. 11 — Ci
Fi
Circulation
l ti wingspan
i
distribution,
di t ib ti
angle of attack = 8deg
bution for winglet equipped wing was verified. In the near future,
more detailed analysis of loads for sideslip flight regimes will be also
carried out.
References:
[1]
[2]
[3]
[4]
Whitcomb, R.T.: A Design Approach and Selected Wind-Tunnel
Result at High Subsonic Speed for Wing-Tip Mounted Winglets;
NASA TN D-8260, July 1976
Hlinka, J., Vanek, F.: Glauert III; Institute of Aerospace Engineering, VUT Brno, 2000
Heyson, H, Riebe, G., Fulton, C.: Theoretical Parametric Study of
the Relative Advantages of Winglets and Wing-Tip Extensions;
NASA TP-1020, 1977
Vanek, F.: An Analysis of the Aircraft Loads Calculation for a
[5]
[6]
[7]
[8]
Computer Automatization and Realization Using Modern HW and
SW Tools; PhD Thesis, 2007
Masak, P.C.: Design of Winglets for Sailplanes; Soaring, June
1993, pp. 21-27
Holmes, B., Dam, C., Brown, P., Deal, P.: Flight Evaluation of the
Effect of Winglets on Performance and Hanling Qualites of a Single-Engine General Aviation Airplane; NASA TM-81892, 1980
Aerodynamic Principles of Winglets, ESDU Report No. 98013,
1998
Popela R., Zikmund P.: Aerodynamický návrh wingletu pro letoun
VUT 001 Marabu; LÚ, VUT vBrně, 2007, LU44-2007-001.AD
Exchanger Integral Property Computation
and CFD Simulation of an Unconventional
Cooling System
Výpočet integrálních charakteristik nekonvenčního chladicího systému
Ing.
g Erik Ritschl / Department
p
of Aerospace,
p
, Czech Technical Universityy
of Prague
Jian Chen, M.Sc. / Concordia University, Montreal, Canada
The paper is focused on the investigation of the integral property of an exchanger. It is important to know these values
for correct setting of the simplified CFD. Another way is to model all details of the exchanger within all other geometry, but it is very time consuming and often impossible to carry out. These integral properties are then used for setting
the radiator boundary condition by means of Fluent solver.
Článek popisuje výpočet integrálních charakteristik deskového chladiče. Jde o výpočet výkonu chladiče a oteplení
proudu vzduchu a tlakové ztráty za chladičem prostřednictvím analytických vztahů a s použitím metod CFD.
Vypočtené veličiny jsou následně použity pro výpočet okrajových podmínek nutných pro výpočet s použitím zjednodušeného modelu chladiče v CFD řešiči Fluent.
Keywords: Exchangers, Cooling system, CFD, Fluent, Cold Jet.
Introduction
The project of a cold jet is a project of a small aircraft powered by fan
inside the fuselage driven by piston engine. The proposed engine is the
Yamaha R-1motorcycle engine. It is necessary to cool this engine.
The project supposes to use an original exchanger, but its accommodation in the aircraft is completely different from the motorcycle. The
exchanger should be placed in the by-pass duct behind the rotor stage
inside the nozzle.
The question is: will the exchanger be sufficient enough or not?
The answer to this question is made by CFD solution in the software
Fluent. It is necessary to find the integral property of the exchanger
such as pressure drop and heat flux. These values were solved by two
different methods, empirical and by CFD.
Description of an exchanger
The exchanger is tabular and originally is placed just behind front
wheel of the motorcycle. To find out some of its data, the power of
a Yamaha motorcycle has been measured. The measurement was
made by Klimeš company. Measured data for one time step are mentioned in Tab. 1.
The average peak power of the exchanger measured during the loading test and valid for engine RPM 10 000 is P cool = 60 KW.
Now one can solve the balance on the exchanger. The amount of
56
C Z E C H A E R O S PA C E P R O C E E D I N G S
Tab.1 — Measured data for Yamaha engine, loading test provided by
Klimeš company
the air is described by equations (1) and (2), where ΔT is from Table
1. This is local minimum of measured value.
M = Q cool*cp*ΔT
(1)
M = 60 000*1006.24*23.2
M = 2.57 kg/s
V = M/ρ
(2)
V = 2.098 m3/s
Exchanger overall geometry:
width
0.338 m
height
0.278 m
heat exchanging area 3.3 m2
front area
0.09 m2
The velocity of the air on the exchanger face is defined by (3):
w = V/ front area
(3)
w = 23.3 m/s
Empirical computation
Pressure loss computation is made for all fluid conditions, laminar,
transient and naturally turbulent sheers flow.
Slat geometry:
height
8 mm
width
1.2 mm
length
26 mm
number of slats
in one row
300
number of rows
26
aspect ratio
0.125
The periodically repeated geometry of the slat is depicted in the Figure 1.
Where f is defined for typical flows by:
Laminar flow: f*Re = 20.585
(5)
Transient flow: f = A+B*Re-1/m
(6)
where
A= 0,0054; B = 2,3*10-8; m = -2/3
Turbulent flow f = 0,0791*Re-0,25*91,0875-0,1125*α
(7)
Solved values are mentioned in Table 2, because the exchanger is placed in the duct behind the rotor we can suppose the turbulent flow.
Value of the pressure loss is then 203 Pa.
Tab. 2 — Exchanger
g pressure
p
loss for basic types
yp of sheer flow
Similarly is solved the heat transfer coefficient h., conclusions are
mentioned in the Table 3.
Tab. 3 — Heat transfer coefficient
Tab
CFD computation
Second method which was applied to this investigation is the CFD
computation by software Fluent. For this purpose was prepared the
computation volume. The exchanger consists of the periodically repeated slats, fig.1. This basic element is chosen as a control volume,
naturally its length is prolonged forward 10 times and backward 30
times. The drawing of the volume is depicted in Figure 2.
Fig. 2 — Domain for CFD computation
Fig. 1 — One half of the exchanger
slat geometry
Subsidiary values:
Hydraulic diameter
based on the slat
geometry
Dh =0.0017 m
Mass flux per unit
front area
G = 27,347 kg/sm2
Reynolds number
based on Dh,
Re= 2836
Definition equations are taken from Fundamentals of Heat Exchanger
Design by Ramesh K. [1].
The value of pressure loss is denoted by:
Δp=(4*f*L*G*G)/(2*?*Dh)
(4)
Computation grid consists of hexahedrons and its dimension is 8x
28x30 for exchanger element. The grid was swept forward and backward and the dimensions are 8x28x50 and 8x28x300 respectively.
This type of grid was chosen to reach value of y+<1, it was obtained
by only one refining during the computation process. The picture of
the grid on the inlet exchanger face is in Figure 3.
Boundary condition was set as wall for S-shaped slat and for the
bounds of the pipe on the top and bottom sides. The temperature of
these walls was set to 400 K. Bottom and upper sides of prolongation was set to periodic boundary condition, sides was set to symmetry.
Fig. 3 — Detail of the grid on the
slat element
57
L E T E C K Ý Z P R AV O D A J
Fluid was modeled as an ideal gas and the viscosity was set to the
Sutherland leading equations. Inlet fluid property was set to gauge
pressure 400 Pa, temperature 300 K and intensity of turbulence 10%.
First computation was done with the standard turbulence k-ε model
and standard wall treatment, number of iteration is 5000, it is a thin
line p tot and p stat in the graph in Figure 4.
Second attempt was made with enabled enhanced wall treatment.
These values are marked by thick solid line. The pressure drop falls
down by 50 Pa. Now the pressure drop is 260 Pa.
Comparison with analytical data shows difference about 60 Pa.
This can be caused by existence of the pipe in the CFD model. Because it wasn´t included in the empirical solution.
It was made a third model where the exchanger was modeled without the slats. The solution is presented in the graph in figure 4 by
dashed curve with subscript simple. The pressure drop is 90Pa.
3/2007
Example of using an exchanger setting
The above mentioned data are used for setting boundary condition
radiator in the Fluent software. As an example is chosen the CFD
computation of the nozzle with a bypass-duct where an exchanger is
located. The geometry of one half of the nozzle and extension of
computation volume in the downstream direction is depicted in Figure 5.
Number of cells in this model with model of heater is 300 000. This
is only one half of the geometry and the boundary layer is sufficient
to achieve value of y+ about 50. So it is necessary to use the standard
wall model. Other settings are listed in Table 5.
Pressure history of the exchanger CFD model
500
400
300
[pa]
200
100
Tab. 5 — Setting of the boundary condition
0
-100
-200
-150
-100
-50
0
50
100
150
200
250
300
Z [m m ]
p tot
p stat
p stat simple
Ptot anhanced wall
Pstat anhanced wall
p tot simple
Fig. 4 — Pressure history, z = 0 -inlet of the exchanger, z = 26 - outlet
After subtraction this ”clean“ pressure drop from complex geometry,
we obtain 170 Pa. This value is under expected 200 Pa.
The computation of the heat transfer:
There are mentioned reported values of the heat flux of one exchanger cell in the table 4. The temperature of all walls was set to 400
K. Names of the zones are pipe which leads the cooling medium, slat
and slat shadow are one and the other side of the slat. So the net heat
transfer for one slat is 6.6 w.
zone 11 (pipe):
zone 12 (slat):
zone 13 (slat -shadow):
net heat-transfer:
0.634652
2.976043
3.051859
6.662554
It is visible from table 7 that the radiators are defined by two leading
equations. The pressure loss is defined by equation (8) [2]. The kL is
non-dimensional coefficient defined by multinominal. Similar is the
definition of heat flux, equation (9). The relation of value h on velocity v is governed by equation (10). The relations of parameters KL and
h are in the graphs in Figure 6.
N
1
Δp = k L ρv 2 , where k L = ∑ rn v n −1 ;0 ≤ N ≤ 7
(8)
2
n =1
q=
m& c p ΔT
A
= h(Tair ,d − Text ), where h = ∑ s n v n −1 ;0 ≤ N ≤ 7
⎡ ρvc p (Tair ,u − Tair ,d )⎤
h=⎢
⎥
⎣⎢ (Tair , d − Text ) ⎦⎥ n
N
(9)
n =1
(10)
[w]
[w]
[w]
[w]
Tab. 4. — Heat Flux for one exchanger slat
Number of slats in the whole exchanger is 300x26x6.66 = 51.9 Kw.
Fig. 6 — The relation of radiator
parameters h and KL
Fig. 5 — The geometry and the computational grid of a nozzle
with bypass-channel
58
C Z E C H A E R O S PA C E P R O C E E D I N G S
This simplification allows analyze the computation by common PC
in acceptable time. The solution was converged after 2500 iterations.
The results are present in the Figure 7 and Table 6.
Tab. 6 — The results of a radiator simulation
Conclusion
The pressure losses given by empirical model are rather lower than by
CFD. The value 170 Pa corresponds to the drop in the middle between transient and turbulent region. This difference can be caused by
complexity of the geometry and its interaction.
The comparison of the heat flux: Input values used in the computation are in Table 7. The label Tw is temperature of the cooling liquid, Tai, Tao are temperatures of incoming and outgoing air, w a is
velocity of incoming air. Compared values are mentioned in Table 8.
Fig. 7 — The Temperature history of bypass
Fig
bypass-duct
duct
Tab. 7 — Input values
Tab. 8 — Comparison of heat flux for the analytical and CFD solution
The main aim was to evaluate the possibility of solving the simple
smallest part of the exchanger to determine its integral properties.
This assumption is correct and hence it is possible to save much computing time and power.
Computing of the pressure loss is in correct rank, but it is not possible to compare CFD data with the analytical solution. The analytical solution operates with a simplified geometry. After the subtraction
of the effect of the geometry complexity the difference is above 15%.
Still the values are in the proposed rank.
Model of the wall treatment is best to use the enhanced wall treatment. The quality of computational grid has to have value of y+<1 [2].
Precision of the heat transfer: Equality of the two mentioned computing methods is precious, relative error is under 0.5% and under 2%
for increasing the flow temperature behind the exchanger.
References:
[1]
[2]
Fundamentals of Heat Exchanger Design, Ramesh K. Shah and
Dušan Sekulic, 2003, ISBN 0-471-32171-0
Fluent user manual v6.1; 2003 Fluent Inc.
Shear Buckling Analysis of Composite
Sandwich Panels
Ztráta stability kompozitových sendvičových panelů při smykovém
namáhání
Ing Martin Baumruk,
Ing.
Baumruk Ing.
Ing Ivan Jeřábek,
Jeřábek Ing.
Ing Karel Barák / Institute of
Aerospace Engineering, Brno University of Technology
The paper deals with the experimental measurement of the buckling critical load of composite sandwich panels under
shear loading and compares the results with theoretical buckling load obtained by calculation. Both Analytical approach
and FEM analysis to calculate the theoretical buckling load are used. Several sandwich panels with different edge
stiffness were tested in the experiment. The critical buckling load was recorded and behavior after the first buckling
mode until the panel destruction was analysed. Different edge stiffness designs were compared. It was observed that the
theoretically calculated buckling loads were slightly conservative in comparison with the actual buckling loads derived
from experiment.
Keywords: Buckling, Composite sandwich panels, Shear load.
Introduction
The use of sandwich composite parts grows in both large general aviation aircraft and small sport aircraft design. (E.g. wing skin, wing
spar web). Composite sandwiches are a layered composite formed by
outer facings (carbon, glass, aramid, or hybrid) and inner thicker lightweight cores (foam, honeycomb). The sandwich provides due to hig-
her effective thickness a good ratio of stiffness to weight. The facings
resist nearly all of the in-plane loads and bending moments and provide bending rigidity. The core forms the construction thickness of the
panel and shifts effectively the facings from neutral axis and also
transmits shear between the facings. The core provides the shear rigidity of the sandwich and supports the facings.
59
The buckling is an important factor in an aircraft design. The buckling resistance can be improved by increasing of the bending stiffness
for example by stringers. The stringers can be manufactured in a relatively simple way for metal structures but for composites construction it means a significant increase in the manufacturing complexity.
That is one of the main reasons why the composite sandwiches have
been massively implemented.
The theoretical assumption is that the panel remains flat up to the
critical buckling load when the system begins to be unstable and panel
buckles out of the plane. The two dimensional skin deformations
occur accordingly to size and edge bounding conditions. After the first
buckling mode the panel is still capable to bear the load but with
changed rigidity and theoretically there can occur other buckling
modes when the load is further increased.In fact the real behavior can
be different and it is strongly dependent on the initial imperfections,
real boundary conditions and other variables. In the case of a sandwich panel with carbon facings there is an assumption that a failure
will probably occur soon after the first buckling mode has occurred
due to the deformation and brittle characteristic of the carbon.
Theoretical modes of failure that may occur in sandwich under
edge load are depicted in Fig. 1. Shear crimping (Fig. 1B) failure
appears to be a local mode of failure, but in fact it is a form of general buckling (Fig. 1A) where the length of wave is very small. Crimping may also occur if the general buckle starts and then the crimp
occurs suddenly because of high local shear stresses and overall
buckle may disappear.
L E T E C K Ý Z P R AV O D A J
3/2007
sured values were used in the buckling calculation. The material characteristic of Divinycell H60 foam was E=70GPa, G=22Mpa, ν=0.3.
The buckling load calculation was done for 24 different panels with
size 350x350mm, 350x700mm, 350x1050mm with different lay-up
(both symmetrical and unsymmetrical), with angles 0° and 45° and
with the core thickness 4mm and 8mm.
Analytical and FEM analyses
At present there are three approaches of critical load buckling calculation in industry: The Analytical and ESDU methods for Shear and
Compression load; Lekhnitskii method for Bending load; and Finite
Element Method Analysis (FEM).
The Analytical (ESDU) approach and the linear FEM analysis are
used in this work.
The Analytical and FEM Buckling Load Calculation methods had
been described in detail in [1] and [2]. The transverse (through the
thickness) shear stiffness was included in the calculation and a method with the neutral plane shift from the middle plane was used for
unsymmetrical panels.
The calculated buckling load Fxycrit under the shear loading
condition for the panel used in the experiment (the fifth type –
no edges modification, the size 350x350mm, symmetric, 4mm
foam, ±45° ply angle, edges clamped) was:
Nxycritt = 42,1 N/mm along the edge
N xycrit = 42,1 ⋅ 350 = 14735 N , Fxycrit = N xycrit + N xycrit
2
2
Fxycrit = 20838N
The FEM analysis was processed in MSC.Patran/Nastran by Buckling, Solution 105. The two models were tested, first with pure 2D
shell elements QUAD4, second with a combination of 2D shell elements QUAD4 and 3D Solid elements for the core model. The frame
was modeled by 1D elements with beam properties and thickness
80mm and height 100mm. The material properties of the frame model
were E = 210000Mpa, ν = 0,3. The beams of the frame were joined
by nodes at the corners with removed rotational degree of freedom
(Pinned DOF). The model was loaded in y axis at corner 3.
The FEM result for the characteristic of the panels used in the
experiment (350x350mm, symmetric, 4mm foam, ±45° ply angle)
and the fifth panel type (no edge modification) modeled by 2D
QUAD4 was Fxycrit = 20522N.
Fig. 1 — Modes of failure that may occur in sandwich under edge load
Specimens and material characteristics
The sandwich panels used in the experiment were symmetric, made
from one layer of carbon biaxial fabric stacked ±45° 200
NLT00 HR1270 0200 Hexcel at the top and bottom and from a 4 mm
thick Divinycell H60 foam. The panels were cured at a room temperature of 22° C for 48hours. Panels with 5 different edge designs
were tested. The edges served for the load transmition from frame to
panel and the relationship between the edges design and the buckling
loads was examined. The first specimen type had the edges stiffened
by an additional layer of the carbon fabric, the second type had the
plywood bonded inside the edges, the third type had the edges filled
with the toughened resin, the fourth type had weaker edges with the
top and bottom layers of the biaxial fabric bonded together alone without the foam core and the fifth type had no modification with original layer of the biaxial fabrics and the foam at the edges.
The size of the experimental panels was 350x350mm. The theoretical, calculated biaxial characteristics were EL = ET =55.8Gpa,
GLT = 3.93Gpa, νLT = 0.0359. The avarage biaxial characteristics from
an experiment performed according the ASTM D3039/D 3039M-00
were EL = ET =36.05Gpa, GLT = 2.67Gpa, νLT = 0.1140. These mea-
Fig. 2 — FEM model
Buckling load — Analytical method
Buckling load — FEM method
20.8kN
20.5kN
The experiment
We developed and used the support and loading frame for the load
transfer to the edges of the panels by flanged joint which simultaneously provided the clamped boundary conditions. The load force and
also the diagonal displacement were measured and recorded. The
deformation of the selected specimens was recorded by strain gauge
roses 1-RY93-6/120. Strain gauges’ temperature compensation by the
help of unloaded specimens with the same properties as the loaded
panels was used. When the critical buckling load was reached the loa-
60
C Z E C H A E R O S PA C E P R O C E E D I N G S
ding was paused to take pictures of the panels. Then the experiment
continued until full destruction of the specimen.
The measured critical load of the fifth panel type (no edge stiffness
modification) was 16.5kN. The lower bearing capacity was due to
a failure at the panel edge in the flanged joint. When the experiment
continued the edges were sheared of by bolts of the flanged joint and
the force decreased. This type of the panel was not further used.
The measured critical loads of the first panel type (the extra stiffened layer of the carbon fabric at the edge) was 26,4kN (specimen
number 3) and 27.1kN (specimen number 8). There appeared shear
failures near the bolts of the flanged joint after disassembling.
The measured critical loads of the second panel type (the plywood
laths bonded inside the edges) was 31,1kN (specimen number 2) and
31,6kN (specimen number 9). The force increased when the experiment continued after the buckling until full destruction of the panel.
No significant shear failures near the bolts of the flanged joint at the
edges after disassembling were found.
The measured critical loads of the third panel type (the edges filled
with the toughened resin) were 21,1kN (specimen number 4) and
36,7kN (specimen number 7). The force increased when the experiment continued after buckling. Cracks appeared between holes for the
bolts of the flanged joint at the edges after disassembling.
The measured critical loads of the four type (the weaker edge with
Fig. 5 — Deformation of the specimen with stiffened edges
after buckling
Fig. 6 — Specimen with stiffened edges at the end of the test
Fig. 3 — The experiment, support and loading frame
the top and bottom layers bonded together without the foam core)
were 21,2kN (specimen number 5) and 28,3kN (specimen number 6).
The force increased when the experiment continued after the buckling. Out of plane deformation of 2mm was recorded before buckling
occurred.
When the critical buckling loads were reached the vertical failure
occurred with one part of the specimen crimped over the other part.
When the loading continued more vertical crimpling failures occurred
or the specimen was torn across.
Conclusion
Fig. 4 — The loading force-time curve for the specimen with the edges
stiffened by the plywood laths
The panels with stiffened edges by plywood laths had the smallest
measured values scatter. The panels with toughened resin in the edges
61
L E T E C K Ý Z P R AV O D A J
3/2007
References:
Fig. 7 — Deformation of specimen with weaker edges
after buckling
[1]
M. Baumruk: Analytical Methods for the Calculation of Buckling in
Composite Sandwich Panels; Letecký zpravodaj. 2005, Vol. 8, No.
3, pp. 10-13. ISSN 1211-877X
[2]
M. Baumruk, P. Průcha: Analytické metody výpočtu ztráty stability
kompozitních panelů a kompozitních sendvičových panelů
[3]
L. T. Tenek and J. Argyris: Finite Element Analysis for Composite
Structures, Kluwer Academic Publishers, 1998
[4]
Bowdler, H. J.: Solution of real and complex systems of linear
equations; Numerische Mathematik 8, pp. 217-234, 1966
[5]
N. D. Phan and J. N.Reddy: Analysis of laminated composite plates
using a higher-order shears deformation theory; Int. J. Numer.
Meth. Eng., 1985
[6]
A. K Noor: Stability of multilayred composite plates; Fibre Sci.
Technol., 1985
[7]
M. Ahmer Wadee and A. Blackmore: Delamination from localized
instabilities in compression sandwich panels; a Department of Civil
& Environmental Engineering, Imperial College of Science, Technology & Medicine, London SW7 2BU, UK, 2000
[8]
Structural Sandwich Composites, MIL-HDBK-23, U.S. Department of Defense, Washington, DC, 1968
[9]
Pierre Minguet, John Dugundji: Buckling and Failure of Sandwich
Plates with Graphite-Epoxy Faces and Various Cores; Massachusetts Institute of Technology, Cambridge
[10]
ESDU 80023: Buckling of rectangular specially orthotropic plates;
ESDU 81047: Buckling of flat rectangular plates (isotropic, orthotropic and laminated composite plates and sandwich panels);
ESDU 89013: Transverse (through-the-thickness) shear stiffnesses
of fibre reinforced composite laminated plates
had bigger values scatter and the manufacturing of these panels is
more complicated. The panels with no edge modification are not
appropriate for this kind of testing. The panels with weaker edges had
lower buckling load and higher out of plane deformation.
The theoretically calculated buckling load by analytical and FEM
analyses was done for the panel with no edge modification. Generally it can be said that the calculated buckling load is lower then the
measured buckling load and the Analytical and FEM methods give
slightly conservative results.
Fig. 8 — Specimen with weaker edges at the end of the test
Fig. 9 The critical buckling load and panel
destruction load
62
C Z E C H A E R O S PA C E P R O C E E D I N G S
Modelling of the Microaccelerometer MAC
Translation Control — Measurement
Channel
Modelování translačního kanálu řízení (měření)
mikroakcelerometru MAC
Ing. Viktor Fedosov / VZLÚ, Plc., Prague
A iis well known, a spacecraft's
As
f ' trajectory
j
on a low Earth orbit
i iis iinfluenced
f
besides
i
the gravitational
i i
fforce by
a number of perturbing factors. It is possible to divide these perturbations into two basic groups: gravitational — due
to Earth's nonideal spheric shape, other celestial bodies attraction (the Sun, the Moon) and non-gravitational — due
to mainly atmospheric drag and direct/indirect solar radiation pressure. A very sensitive electrostatic feed-back threeaxes and symmetrical microaccelerometer MAC with the cubic proof mass has been developed for the measurement
of accelerations arising by virtue of non-gravitational perturbations acting on the spacecraft board. The article
describes a performance model of the microaccelerometer measurement channel in one axis and its analysis.
Článek popisuje postup simulace funkce translačního kanálu řízení mikroakcelerometru MAC v prostředí SIMULINK.
Základem modelu jsou bloková schémata kanálu a systém rovnic popisující funkční vazbu mezi vstupem (zrychlení)
a výstupem (napětí). Pozornost byla věnována simulaci polohového detektoru, který je charakterizován koeficientem
přenosu, vlastním šumem a teplotní závislostí (variace nuly). Podmínky modelování odpovídají realizačnímu projektu
TEASER — rozměry a hmotnost družice, parametry oběžné dráhy, teplotní prostředí.
Keywords: non-gravitational effects, non-conservative accelerations, Microaccelerometer, real-time
measurement on spacecraft board.
Principle of microaccelerometer measurement
The microaccelerometer's sensor is composed of a cubic proof mass
free flowing in the cubic cavity. Proof mass is produced from quartz
glass and the cavity is composed of six prisms from the same material. The centre of the sensor should be placed in the spacecraft's centre of gravity. Proof mass is separated from external influences by
satellite structure and microaccelerometer construction. Free motion
of the proof mass is realized by virtue of only gravitational law. The
cavity is rigidly connected to the satellite body. Gravitational and also
all other perturbing forces acting on a satellite produce its acceleration and it is the same as the cavity acceleration. The difference between the acceleration of the cavity and the acceleration of the proof mass
is the sum of all accelerations produced by non-gravitational forces
acting on a satellite. The proof mass is moved to the centre of the
cavity by an electrostatic feed back servomechanism. In such case the
electrostatic servomechanism must generate force to the proof mass
which is proportional to the sum of all non-gravitational accelerations
of the spacecraft. Accordingly, input signal to accelerometer measurement channel is non-gravitational acceleration (m/s-2 in case of linear disturbance effect or rad/s-2 in case of angular disturbance effects)
and output is voltage (V).
Model description
A block diagram of the proof mass position feedback control system is shown in Figure 1. This scheme is described
by following dependences (system of equations):
(1)
Ux = KD H1 (s) x
Γsum - Γκ
Γκ) H(s)
x = (Γ
Γκ = ((x/D) - (Ux/Uo))(A1
Γκ
( *U0^2)
Where:
Γsum - input summarized acceleration
Γκ
Γκ
- proof mass compensative acceleration
Ux
- regulator compensative voltage
m
- mass of proof mass A1 = 4(C/D)
D
- distance between electrode and the proof mass in the
central position
C
- capacitance between one electrode and the proof mass
Uo
- polarization voltage
H(s) - 1/s2- transfer of acceleration to proof mass shift
- POSDET Sensitivity
KD
H1 (s) - regulator transfer function
PD
regulator with following parameters is used in the microaccelerometer servo system:
Gain coefficient (Kp) = 8;
Time constants T1 = 0.332 (s), T2 = 6.64 (ms)
Regulator Transfer Function H1(s) is described following expression:
H 1 ( s ) = Kp
(1 + T1 s )
(1 + 0.332s )
=8
(1 + T2 s )
(1 + (6.64 E − 3) s )
H2 (z) - ADC transfer function
”code“ - accelerometer outputs in digital format
It necessary to point out that the use of common control constants:
A1, A2, A3 allows us to apply the scheme shown in Fig. 1 (and its
model, Fig. 2) to simulation of the rotation channel. Difference among
these constants magnitudes for translation and rotation channels are
described in [3]. In this case the basic system of equations can be
63
described by the following expressions:
Γκ = A1 U02 (A
Γκ
(2)
( 2x - Ui/Uo)
- proof-mass compensative accelerations;
Ui = A3 H1(s) x
- regulator compensative voltage;
x = (Γκi - Γsumi)
- linear (angular) shift of proof-mass.
Consequently:
- PD Regulator transfer function:
H1(s) = Kp(1+t1s)/(1+t2s) = 8*(1+0.332s)/(1+(6.64e-3)s);
- Open loop transfer function:
H0(s) = [A
[ 1KdKp(1+t1s)]/[(1+tt2s)(-A1A2+s2)];
- Transfer:
[U/Gk]=[KdKp(1+t1s)]/[A1KdKp(1+t1s) + (1+t2s)(-A1A2+s2)];
Figure 2 shows the model in Simulink of the one axis translation
channel. This model consists of the following connected functional
elements which allow us to simulate and analyze basic performances
of the circuit:
Simulation of proof mass drift. This part of model consists of ”Transfer X'' to X block“ — double integration of acceleration (the result of
this operation is proof mass displacement from base point on accelerometer coordinate axis) and Saturation block. This block simulates
contact between proof mass and wall-stop. In case the proof mass is
touching the wall-stop, output signal from this block is equal to upper
or lower limit (+2e-5 m).
Simulation of Position Detector (POSDET). This model part reflects the following Position Detector parameters: POSDET Sensitivity — KD (as Gain POSDET A3 block), POSDET noise characterized
by power noise and its sample period and POSDET temperature zero
variation. Temperature influence on POSDET output is simulated by
special Temperature File which will be described hereinafter in
”Modeling of inputs“.
Simulation of CONTROL which is presented by Regulator and AD
Converter Subsystem. Summarized signal from POSDET is input signal to Regulator. Regulator is modeled by Gain Kp — regulator gain
and regulator transfer function (Transfer Fcn1 Block). Regulator output signal input to anti-aliasing filter and AD Converter block — Output signal from AD Converter subsystem is accelerometer output. Ux
— compensative regulator voltage is transformed to proof mass compensative acceleration by Proof Mass Dynamic simulation part —
it presents the feedback of control (see Fig. 2). In the frame of the
paper we will not envisage the problems related with analog-digital
converting and we are limited to translation channel analysis.
L E T E C K Ý Z P R AV O D A J
3/2007
Definition of the model inputs.
Before an examination of the model performance, some remarks
about modeling conditions and man-made assumptions are desirable.
1. We will examine one translation control circuit facing in direction of oppositely to spacecraft velocity (in direction of perturbation
called the drag). Simulation time is equal to one orbit. Suppose, that
a spacecraft which is characterized by size (0.5 x 0.5 x 0.7 [m]) and
mass (70 kg) moves at orbit with following parameters:
orbit period — 100 minutes;
circular orbit;
altitude — 450 km;
inclination — 98,8°;
perigee argument — 66,5°;
eccentricity — 0,00112.
Accordingly, drag force (R) is possible to calculate as
1
R = C D ρAV 2, here:
2
CD is aerodynamic drag coefficient, for the simulation case = 2.2
A is effective cross-sectional area = 1/4 of the spacecraft surface area
ρ is thermosphere (high atmosphere density), for the simulation case
is calculated by using of the thermosphere model TD88 in relation
to reference environmental condition (solar flux and geomagnetic
parameter) and spacecraft orbital parameters. TD88 detailed description is given in [1] and [2].
V is spacecraft velocity relative to atmosphere rotation, for the simulation case it is possible to define high atmosphere as statically and
V is equal to the spacecraft absolute velocity.
Spacecraft acceleration (aD) due to drag force is following:
aD = R/m
/ , here R is drag force, m is S/C mass.
Values of the atmosphere density and S/C accelerations due to drag
disturbance force for simulation case are shown in diagram hereinafter (Graph 1):
Graph 1 — Atmosphere density and drag
64
C Z E C H A E R O S PA C E P R O C E E D I N G S
2. The most crucial part of the microaccelerometer is POSDETposition detector of proof mass. This element is characterized by sensitivity (100 000 V/m), basic noise and zero variation due to temperature influence. According to [4] spectral density of the POSDET noise
has character of white noise in frequency range from 0.001 Hz to 1
Hz. The smoothed noise PSD of position sensor is plotted in Graph 2
(expressed in ms-2/sqr Hz). Time diagram during one orbit of POSDET noise expressed in V is shown in Graph 3.
Graph 2.
POSDET
Noise PSD in
[ms-2/sqrHz[
Where:
B0 = -2.27E-7/-3.03E-10 (ms-2)
is Bias of the measured channel;
S = 3.39E-5/4.43E-5 (V/ms-2)
is sensitivity of the measured channel
Measurement uncertainty for first case is 3.7E-9 (ms-2) (Graph 7), for
second case is 6.84E-11 (ms-2) (Graph 8)
For estimation of each models adequacy we use Residual analysis
(by residual p-graph) and the Determination Coefficient R2 (which
indicates ratio of analyzed variables variability, for example if
R2=0.99 that model explains 99% variability of dependent variable).
General time diagram for Drag and Microaccelerometer Outputs is
presented by Graph 9.
Obtained equations show the main problem related to microaccelerometer MAC operation during orbital measurement is requirement to
high accuracy calibration in orbit (including POSDET thermal dependence definition) or minimization of the thermal cycling effects.
Graph 3.
POSDET Noise
Time diagram
during one orbit
Graph 5 — Drag versus U_out including POSDET
thermal zero variation
POSDET temperature zero variation is modelled on base of POSDET
thermal-vacuum test results. Detailed description of the test procedure and obtained results are given in [5]. For this simulation case we
suppose that temperature environment at POSDET corners matches
to values defined during tests and range of POSDET zero variation
matches to measured data (see the Graph 4).
Graph 6 — Drag versus U_out ”cleaned“ from thermal influence
Graph 4 — Thermal and POSDET zero variations during one orbit
Results of simulation.
We obtained following linear regression equations which describe the
interrelation between Drag (inputs to microaccelerometer) and measured data (U_out — microaccelerometer outputs):
Drag = -2.27E-7 - 3.39E-5*U_Out
including temperature zero variation (Graph 5)
Drag = -3.03E-10 - 4.43E-5*U_Out
”cleaned“ from temperature influence (Graph 6)
Graph 7 — Comparison among Drag and measured values (including
POSDET thermal zero variation)
65
L E T E C K Ý Z P R AV O D A J
Graph 8 — Comparison among Drag and measured values
(”cleaned“ of POSDET thermal zero variation)
References:
[1]
[2]
RNDr. Ladislav Sehnal DrSc., RNDr. Libuše Pospíšilová:
Výpočetní simulace negravitačních zrychlení pro podmínky
projektu TEASER (č. OV6050097 Astronomický ústav AV
ČR)
L. Sehnal. L. Pospishilova: Thermospheric Model TD88;
Astronomical Institute, Czechoslovak Academy of Sciences. Preprint N67
3/2007
Graph 9 — Drag and related Microaccelerometer
Outputs time diagram
[3]
[4]
[5]
Milan Chvojka, Viktor Fedosov: Microaccelerometer MAC
Control Concept Analysis; Proceedings of 14th Saint Petersburg International Conference on Integrated Navigation
Systems, ISBN 978-5-900780-67-2
M. Chvojka: ACC Performance Analysis, SW-AN-VZLAC-003, Issue 2.5
POSDET Thermo-vacuum Test Report; Doc No.: SW-TNVZL-AC-0009, Issue 0.1
Impact Analysis of the Rigid Body on the Thinwalled Aluminum Structure with Considering of
the Stochastic Material Properties
Simulace nárazu tuhé trubky na tenkostěnnou duralovou
konstrukci s experimentem
Michal Mališ / Institute of Aerospace Engineering, Brno University of
Technology
The article compares a simulation of a rigid tube impact on an aluminum structure with an experiment. The simulation
was performed
f
d with
i h nonlinear
li
explicit
li i transient
i
code
d MSC
MSC. D
Dytran. The
Th experiments
i
were carried
i d out iin A
Aerospace
Institute laboratory BUT. An additional analysis of the influence of stochastic material properties on simulation was
done.
Článek popisuje porovnání simulace nárazu tuhé trubky na tenkostěnnou duralovou konstrukci s experimentem.
Numerická simulace metodou konečných prvků (MKP) byla provedena v systému MSC. Patran/Dytran. Experimentální
část práce i stavba zkušebního zařízení byla realizována na zkušebně Leteckého ústavu VUT v Brně. Práce je doplněna
o analýzu vlivu stochastických vlastností duralového plechu na výsledky MKP simulace.
Drop tower building
Introduction
Finite elements methods (FEM) are widely used effective methods for
reaching required aircraft structure properties. This is the only tool,
expect crash test, for crashworthiness application how to predict behavior of structure.
Recently we can see strong effort to engage maximum parameters
to FEM simulation which affect the results. A lot of them have stochastic character (material property, angle of impact, impact velocity
act.).
Present work is divided to several stages. Building of experimental
equipment and test of measuring devises. Carrying out experiment rigid tube impact to thin-walled structure. Tuning of FEM model and
verifying with experiment. The last stage describes probabilistic analysis stochastic material property influence on FEM simulation
results.
The drop tower was designed for impact tests of specimens and small
structure assemblies (Figures 1 and 2). It is quickly flat pack and utilizes a lot of laboratory equipment. The tower consists of two leading
rods (diameter 40 mm) with a moveable dropping weight (20,8kg).
The leading rods are standing between the vertical stands. The dropping weight is manually lifted to the appropriate height and drop
weight is released using a remote control.
The kinematical values (displacement, velocity and acceleration) of
the dropping weight during the impact are measured with high-speed
camera (Figure 3).
Experiment
The thin-walled structure for impact test was made from the 1.5 mm
thick 2024 aluminum alloy sheet plate. The structure consists of two
parts riveted with 38 rivets (rivet diameter 3,5mm). Figure 5 shows
the specimen structure.
66
C Z E C H A E R O S PA C E P R O C E E D I N G S
Fig. 5 — Thin-walled aluminum specimen
Fig. 1 — Drop tower plan with
basic dimensions
Fig. 2 — Drop tower
contact was applied between the specimen and impact tube with the
static friction coefficient (fs=0.17), dynamic friction coefficient
(fk=0.15) and contact thickness 1. The initial velocity of the impact
tube was 6,45 m/s.
The bilinear material model and BLT (Belytschko-Lin-Tsay) property element formulation was used for sheet elements. The impact
tube property was defined by RIGIG material model and default set
for PSHELL definition.
Figures 8 to 10 show comparison between a computed displacement velocity and acceleration results with the impact test data. The
black curves represent impact test results data and the color ones are
simulation results from three model types. The models differed in global element size (10 mm, 5mm and 3mm). Figures 12 to 15 show
comparison between high-speed film data with the model.
Stochastic material property analysis
Fig. 3 — High-speed cameras
Fig. 4 — Drop weight
All three impact tests were carried out from 2,15m height with
impact velocity 6.45 m/s. The specimen before and after execution are
shown on the Figures 6 and 7.
During the test specimens stand on two support structures with
390 mm distance between them. The specimen was fixed to the support structure using two clips.
The dropping weight tube was skewed to symmetry under angle
26° (Figure 4) so that the impact could be easily visible by high-speed
cameras. The exact position of the measuring target located inside the
dropping weight tube was recorded by high-speed camera. The evaluation of the kinematical values was performed with Matlab software package. See the impact test evaluation on Figures 8, 9, 10.
The influence of stochastic material properties on numerical simulation was analyzed using Response Surface Methodology (RSM). The
method is suitable for solution of complex problems with reciprocally independent design variables. Block diagram on Figure 11 shows
the analysis procedure.
Tension modulus yield stress and hardening modulus were chosen as
independent variables. The tension test was used for material data
assessment. The mean values were figured from tension test. The boundary values were approximated (Table 1). If real values from tension test
were used the results would be very close. In addition to this the RSM
analysis looses the accuracy close to the boundary values therefore it is
better to enlarge the boundary interval. The boundary values were calculated by mean values and coefficient multiplication (0,9 and 1,1 for
modulus and 0,8 and 1,2 for hardening modulus and yield stress).
The Box-Behnken design of experiment was used. The 13 numerical simulations were necessary for 3 variables RSM analysis.
Tab. 1
The model with the sparsest mesh (element size 10mm) was chosen for
RSM analysis purpose. The elapse time of a simulation was 12,3 minute.
The results of the Monte Carlo analysis are distribution function of
the maximum displacement probability and maximum deceleration
Simulation
The numerical simulation was performed using nonlinear explicit transient dynamic code MCS. Dytran. It is suitable for simulation of
a large deformation, crashworthiness, bird impact simulation, initiation of an airbag and its interaction with passenger as well as explosion
in a container.
The crash test simulation model was consisted of aluminum specimen and impact tube. 2-D square QUAD4 elements were applied on
the both parts of the model. The bottom side of the specimen had constrained displacement in all six degrees of freedom. The Master-Slave
Fig. 6 — The structure
before impact
Fig. 7 — The structure
after impact
67
L E T E C K Ý Z P R AV O D A J
Figuree 8 — Displacement simulation results compared with experi
exper ment
3/2007
Figure
gure 9 — Velocity simulation results compared with experiment
experim
Figure 10 — Acceleration simulation results compared with experiment
expe
The stochastic material property analysis slightly exceeds effective
usage for present application. The real differences of calculation caused by stochastic material property are, for aluminum structure, lower
than inaccuracy of simulation. Therefore we should engage in better
tuning of simulation. The stochastic material property analysis could
be important for material with higher property dispersion. The probabilistic analysis could be useful tools for analyzing of widely used
composite materials structures.
Figure 12 — Simulation versus
experiment in time t = 0.005 s
Figure 11 — RSM procedure
probability of the impact tube during the crash. Other figures (not included here) show the distribution functions of the results in comparison with distribution function of normal distribution.
Results
The simulation results of the displacement and velocity correlate with
experiment sufficiently. Inaccuracies are in 5%. Acceleration results
accuracy are in interval 13-24%. Technical applications usually require displacement and deformation results data. Model mesh density
influences acceleration results by comparison with displacement and
velocity results. Elapse simulation time between the model with dense
mesh and sparse mesh increased 16 times.
Figure 13 — Simulation versus
experiment in time t = 0.010 s
Conclusion
The simulation using MSC Dytran application verified sufficient correlation with experiment. Similarly the bilinear material was succeeding for present crash analysis application.
Figure 14 — Simulation versus
experiment in time t = 0.012 s
References:
Todoki A.: Response Surface Methodology; Tokyo Institute of Technology, 2003
[2]
Myers R. H., Montgomery D. C.: Response Surface Methodology:
Process and Product Optimization Using Design Experiments; John
Wiley&Sons Inc., 1995
[3]
FAA: Probabilistic Design Methodology for Composite Aircraft
Structures, Virginia, 1999
[ 4 ] MSC. Dytran User's Manual Version 2004; MSC. Software Corporation, Los Angeles, CA, 2004
[1]
Figure 15 — Simulation versus
experiment in time t = 0.021 s
68
C Z E C H A E R O S PA C E P R O C E E D I N G S
Czech Aerospace Research Centre — Workshop 2007: / Seminář CLKV 2007
Addendum:
Directory of Papers not Published in this Issue
Seznam nepublikovaných příspěvků
To bring a complete overview of this
year's proceedings the following list sums up all
other papers presented at the 2007 conference
that could not be published in this issue due to
capacity restriction.
Ing. Armand Drábek, Ing. Zbyněk Hrnčíř, Ph.D.
(drabe
r [email protected]
z z ; [email protected]
i
z z):
Potential of Blended-wing-body Aircraft in Future Air
Transport
Možnosti letadel typu samokřídlo v letecké dopravě
Ing. Miroslav Šmíd, Ing. Vítězslav Hanzal ([email protected]
i
z z;
[email protected]
z
z z):
Flow Simulation in Cleaning Element of Spining Unit
Simulace proudění v čistící vložce spřádací jednotky
Klínek P. (kline
i [email protected]
z z):
Aeroacoustic Analysis of the Propeller by AAV_M3A7 and
L16_H10A1 Software
Aeroakustická analýza vrtule pomocí programů AAV_M3A7
a L16_H10A1
Ing. Jan Tůma ([email protected]
z z):
Verification of Using Steady State or Unsteady Model in
Optimization Process of Blade Passage
Porovnání stacionárního a nestacionárního modelu radiálního
kompresoru
Ing. Tomáš Jamróz, Ing. Jaromír Lamka, CSc., Ing. Karel
z z; [email protected]
z z; [email protected]
z z):
Patočka (pato
(p [email protected]
Combustion Turbine Durability Calculations
Výpočet životnosti spalovací turbiny
Ing. Milan Merkl, CSc. ([email protected]
z z):
Some “Marginal Tasks” of Dependability Solution
Některé "okrajové úlohy" řešení spolehlivosti
Ing. Finda Jindřich. (f
(fi
[email protected]):
Reability Centered Maintenance (RCM)
Údržba systému (soustav) letadel zaměřená na bezporuchovost
f .vutbr.cz
r z):
Šplíchal J. (spli
p [email protected]
Human Factor in Civil Aviation
Lidský faktor v civilním letectví
Posters / postery:
Ing. Petr Bělský ([email protected]
ky z z):
New FSW Equipment in VZLÚ, Plc.
Nové vybavení pro třecí svařování ve VZLÚ, a.s.
z z):
Miroslava Nová ([email protected]
Dynamic Fault Tree Implementation for Assessment of
Human Factors Impacts
Uplatnění dynamických stromů poruchových stavů při hodnocení vlivu lidského činitele
Room for Your Notes
Numerical Study of Steady and
Unsteady Flow in a Centrifugal
compressor
Colo r illustrations
Colour
ill strations to the article published
p blished on
pages 38-40.
Fig. 5 — Relative pathlines based on tip gap of blade
Fig. 1 — Pathline through impeller and diffuser solved by CFX with
showed grid
Fig. 6 — Relative pathlines based on 50% height of the blade
Measurement and Evaluation of the Interior Noise in Aircraft
Colour illustrations to the article published on pages 22-27.
Fig. 13 — Examples of an 3D FFT spectrogram of the interior noise in
a turboprop aircraft during its climbing
Fig. 14 — An example of another 3D representation of an interior
noise spectra in a turboprop aircraft during its start to take-off
Fig. 15 — An example of SPAD analysis of the interior noise in a turbopropeller aircraft; the analyzed 3-minute time segment is demarked by
a rectangle in Fig. 8
Contents / Obsah
Aerodynamic Design of V48 Model Propelleers
Aer
Ae
erod
ody
dyn
yna
nam
ami
m cký
ký ná
náv
ávr
vrh
h mo
mod
odelov
ový
výc
ých
ch vrt
r tu
ulíí V4
ul
V48
48
Measurement of Tension and Torq
que Moment of
Propeller
Měře
ření ta
tahu a kr
krout
uticcí
cíh
ho mo
ho
mom
ome
men
entu
tu
vrttulee
vr
Appliccatio
on of Artifficcial Neural Networks for
Searchin
ng of Faultty Components of Turbin
ne
Engin
ne
Evaluatio
on Methodology of RTD Projeects —
Chapter Verifficcatio
on of Evalu
uation Methodolo
ogy
at Specific Projeects
Po
P
ouž
ou
uži
ž tí
tí neuro
ron
ono
nov
ovýc
ých
ch sí
síttíí k na
nal
a ez
eze
z en
ení
ní po
poš
oškoz
oze
z en
ené
né
části tu
čá
turb
rbi
bin
nov
no
ového mo
mot
o toru
ru
Ho
H
odn
od
dno
not
o ticcí
cí me
met
eto
odi
od
d ka pro
rojekt
ktů
ů vý
výz
ýzkumu
mu
a výv
ývo
voj
ojee — ka
kap
api
p to
tol
olaa Ver
erifika
kac
ace
ce modelů na
na
kon
ko
onkr
kréét
étn
nícch pro
ní
roj
o ekt
ktec
ech
ch
Analyysiss of Fatig
gue Crack Growth Under the
Spectrum Loading
Výýp
V
ýpo
poč
oče
čet šíř
ířeen
ení tr
trh
hliin
hl
ny př
ny
přii za
za tě
těžo
žov
ová
vá n í
spe
sp
pekt
ktre
re m
Measurement and Evalu
uatio
on of the Interior
Noise in Aircraft
Měře
ření a hodno
noc
oce
cen
ení
ní vn
vnitř
třn
níh
ní
ho hl
ho
hlu
uku
uk
ku
v le
leta
tadlec
ech
ch
Compositte Recycling — Technology, Recyclees,
Theirr Paraameters and Possible Appliccatio
ons
New FSW Equipment at VZLÚ, Plcc.
Buckliing of Shellls
Numerical Study of Steeady and Unsteady Flo
ow
in
n a Centriffugal Compreessor
Industriaal Measurementss of Frequency
Characteristiccs of Small Sport Aircraft
Optimizzation of Stifffened Panel — Desig
gn of
Testing Equip
pment
Design and Manufactu
ure of Airrcraft Parts
by LF technolo
ogy
Ducted Fan Power Unitt Demonstrator
for Ulttra Light Airplaanes
The Methodolo
ogy of Wingleet Aero
odynamicc
Desig
gn for UL and VLA Aircraft
Exchanger Integral Property Computatio
on and
CFD Simulaatio
on of an Unconventional Cooliing
System
Shear Bucklin
ng Analysiss of Composite
Sandwicch Panels
Modellin
ng of the Miccroacceleero
ometer MAC
Translaation Control — Measurement Channel
Impact Analyysis of the Rigid Body on the Thinwalleed Alumin
num Structure with Consid
dering of
the Stochasticc Materiaal Properttiees
Reec
R
ecy
cyk
ykl
k ac
ace
ce ko
kom
ompozi
z ittů - te
tech
chn
hno
nol
olo
ogi
og
g iee,
e , re
rec
ecykl
kláát
átyy,
y
jej
je
e ich pa
par
aram
ame
met
etrryy a mo
mož
ožné ap
apl
pliikka
kac
ac e
No
N
ové
ov
vé vy
vybav
ave
ven
ení
ní pr
pro
o tř
tře
řecí
cí sv
sva
vařo
řov
ová
vání ve
ve
VZL
VZ
ZLÚ
LÚ, a.
a.ss.
Staab
St
abi
billiittn
ní úl
ní
úlo
ohy
oh
hy skoře
ř e p in
Po
P
oro
or
ovn
ov
vnán
ání
ní st
staacci
cio
onárn
on
rního
ho a nest
staac
acio
c on
oná
nár
árn
níh
ní
ho
ho
řeš
ře
eše
šení od
ods
dst
střřeed
edivvé
véh
ého
ho ko
kom
omp
mpr
prees
esoru
ru
Prrůmy
P
mys
yslov
ová
vá měř
ěře
ření
ní fr
freek
ekv
kve
venč
nčn
ční
nícch
ch cha
har
araak
akt
kteeri
r st
stik
malýýc
ma
ých
ch sport
rto
ovnícch
ov
ch le
leto
t ou n ů
Optimal
aliizza
zace
ce vy
vyz
yzttu
užen
už
ené
néh
ého po
pot
otaah
ahu
hu — ná
návrh
rh
zkuše
zk
šeb
ebn
bní
níh
ho za
ho
zaří
řízz en
ení
ní
Ko
K
ons
on
nst
strrukc
kce
ce a vý
výro
rob
oba
ba so
souč
učá
čás
ást
stíí le
leta
t ad el
tec
te
ech
chnol
o og i í L F
Deem
D
emo
mons
nst
strráát
áto
or ve
or
vent
ntiilláát
áto
oro
rovvé
véh
ého po
poh
oho
hon
onu
nu pro
ro
ulttrraaleeh
ul
ehk
hká
ká le
let
etaad
dlaa
dl
Metodika
ka ae
aer
ero
ody
od
dyn
ynamicck
cké
kého náv
ávr
vrhu wi
win
ngl
ng
gleet
etu
u
pro
pr
o UL a VLA
LA leto
tou
ou n
Výýp
V
ýpo
poč
oče
čet inte
tegrá
rál
á níc
ích ch
cha
har
araak
akt
kteer
eriissti
tk
nek
ne
eko
konv
nve
ven
enč
nčn
ční
níh
ho chlad
ho
adiccí
cíh
ho syst
ho
stém
ému
mu
Zttrrát
Z
átaa st
staab
abi
billiittyy ko
kompoz
ozi
z itto
ový
ov
výc
ých se
sen
end
ndv
dvi
viččo
čový
výc
ých
ch
pan
pa
ane
nelů při
ři smy
myk
yko
kov
ové
vém nam
amá
máh
áhá
hán
ání
Mod
odelo
ová
ov
ván
ání tr
traan
ans
nsl
s ačního
ho ka
kan
aná
nál
á u ří
řízze
z ení
(mě
(m
měř
ěřeen
ení
ní)) mi
mikkr
kro
oak
oa
akce
cel
eleer
ero
ome
om
met
etrru
u MA
MAC
AC
Siim
S
mul
mu
ulaace
ce ná
nár
áraaz
azu
zu tu
tuh
uhé
hé tr
tru
ubk
ub
bky
ky na
na te
tenkos
ost
stěěn
ěnn
nno
nou
ou
dur
du
uraal
alo
ovo
ov
vou ko
kons
nst
strrukc
kci
ci s ex
experi
r men
ent
nteem
em
© C Z E C H A E R O S PAC E
M A N U FAC T U R E R S A S S O C IAT I O N

Podobné dokumenty

czech aerospace - Výzkumný a zkušební letecký ústav

czech aerospace - Výzkumný a zkušební letecký ústav Aeronautical Research and Test Institute / VZLÚ, Plc. Beranových 130, 199 05 Prague 9, Letňany Czech Republic Phone.: +420-225 115 223, Fax: +420-869 20 518

Více