TWENTYFIFTH EUROPEAN ROTORCRAFT FORUM
Paper n• G22
COAXIAL HELICOPTER ROTOR DESIGN & AEROMECHANICS
BY
Boris N. Bourtsev, Sergey V. Selemenev, Victor P.Vagis KAMOV Company, Moscow Region, Russia
SEPTEMBER 14- 16, 1999 ROME ITALY
ASSOCIAZIONE INDUSTRlE PERL' AEROSPAZIO, I SISTEMI E LA DIFESA ASSOCIAZIONE ITALIANA DI AERONAUTICA ED AS TRONAUTICA
G22- I Introduction The tip vortex coaxial rotor wake was visualized in a forward flight. In the rotor front part the free tip vortices were positioned above The advanced technology is based on the tip path planes. That flat part of a free ( mathematical simulation ( software ), on a model testing, on a flight tests and on an wake could extend along the rotor up to 314 of experience of helicopter development, on the its radius ( Fig.5 ). new developed materials. The Fig.!A & Fig.1B show basic 2. Basic design solutions and parameters of coaxial KAMOV's helicopters. aeroelastic phenomena
1. Coaxial rotor mathematical It is very important to have a substantiation simulation & physical experiments of aeromechanical phenomena. This is feasible given adequate simulation making possible to explain and forecast: The survey of the theoretical and - natural frequencies of structures; experimental coaxial rotor aerodynamic - loads & deformation; research published by Coleman, C.P. [4] and - aeroelastic stability limits: stall flutter, by Bourtsev, B.N. [9,10]. transsonic flutter, ground resonance; The paper [9,10] presents the results _of - helicopter performance & maneuverability. figure of merit analysis for single and coroaal KAMOV Company has developed software to maln rotors at hover as well as for a helicopter simulate a coaxial rotor aeroelasticity with a tail rotor and a coaxial helicopter. The [1,2,3,8,9,10]. Aeroelastic phenomena to be analysis has been performed using a simple simulated are shown in the Fig.6 as ( 1-7 ) physical model based on the results of lines in the following way: numerical simulation, wind tunnel tests and (1) -system of equations of coaxial rotor full scale flight tests. blades motion ; It is demonstrated [9,10] that a (2) - elastic model of coaxial rotor characteristic feature of coaxial maln rotor is a control linkage (boundary data); high aerodynamic perfection at hover caused (3) - model of coaxial rotor vortex wake; by an additional amount of air being sucked in (4,5,6) - steady and unsteady aerodynamic by the lower maln rotor ( Fig.2 ). The coaxial airfoils data; rotor at hover demonstrates a 13% larger figure (7) - elastic I mass I geometry data of of merit value in comparison with a single the upper /lower rotor blades and rotor unbalanced by torque ( Fig.2 ). Absence of the hubs. of tail rotor power losses provides a 20% The lines (1-8) in Fig.6 show functional larger figure of merit for a coaxial helicopter capabilities of the software. Columns (1-5) as a rotorcraft ( Fig.2, Fig.4 ). Fig.3 presents conform to versions of the software . Both the coaxial rotor figure of merit. These results steady flight modes and manoeuvres of the were obtalned by measuring in full scale flight helicopter are simulated . test at hover. Based on experience ofKAMOV Company new key design approaches regarding the Full-scale flight investigation of a Ka-32 coaxial rotors of the K.a-50 helicopter were coaxial helicopter tip vortex wake structure developed. was successfully completed [5 ,6]. A smoke The blade aerofoiles were developed by visualization method was applied using blades TsAGI for the K.a-50, K.a-115, Ka-226 smoke generators installed. A tip vortex wake helicopter specially (Fig. 7). Optimal was visualized in hover, at low flight speeds combination of CL,Co,CM( a,M) characteristics and medium flight speeds ( Fig.3, Fig.5 ). was a necessary condition to achieve: A specific dimensionless number has been - high G-load factor & stall limit; adopted to determine inductive velocities and - acceptable margin of flutter speed ; flight speeds related to an ideal inductive - low loads of rotor & linkage ; velocity in hover ( Fig.5 ). The wake form was - low vibration level ; determined and the tip vortex velocities were - high helicopter performance. measured in hover. The tip vortex vertical Blade sweep tips developed by KAMOV velocities turned out to be less then the Company affords for the same purposes. inductive velocity of the ideal rotor in hover. Usage of all key approaches regarding ( The measured wake contractions were equal to rotors becomes an sufficient condition to 0.85R for the upper rotor and 0.91R for the achieve high performance of the rotor system lower rotor ( Fig.3 ). and therefore of the helicopter as whole.
G22- 2 3. KAMOV's Key Technologies for the 11 approximation-calculation" of the matrix-function elements. With the help of this formula the rigidity characteristics of the 3.1 Fiberglass & fiber graphite control linkage aggregates were determined rotor blades without their physical measuring for coaxial helicopters of four types. It is demonstrated In the end - 1950's KAMOV Company that the eigenvectors of the elasticity matrix designed, built and tested fiberglass rotor are actually torsional modes of the six blades blades. In 1965 first production fiberglass rotor on the control linkage and the eigenvalues are blades were successfully flown on the actually linkage "dynamic elasticities" of the KAMOV Ka-15 helicopter. In 1967 first corresponding mode which are usually production fiberglass rotor blades were measured in a different way - that is by the successfully flown on the KAMOV Ka-26 linkage frequency testing. The study results are helicopter. In the end 1970's the graphite & illustrated onFig.9 [!]. glass fiber rotor blades were successfully produced by KAMOV Company. 4. The basic design aeromechanic The graphite fibers had potential specific stiffuess values which six thnes those of & aeroelastic problems of coaxial rotor current aircraft materials. This proved to be helicopter have been developed and highly significant in addressing elastic stability key technologies have been achieved and aeroelastic design problems. The combination of glass and graphite 4.1 KAMOV Company experience fibers resulted in excellent, benign failure modes and provided aerodynamics, structures concentrated in the Ka-50 attack and dynamic engineers. helicopter Ka-50 helicopter advanced rotors geometry is: - special airfoil; Acceptable margin of flutter speed and stall - blade opthnal twist; flutter speed of the Ka-50 helicopter were - blade swept tip ( Fig. 7 ). substantiated by mathematical shnulation and All KAMOV's blades fitted with validated by flight test results ( Fig.! 0 ). Flight electrothermal de-icing system. tests results are shown by ( VrAS- roR) relation in Fig.JO. A part of flight test points is given 3.2 Advanced rotor blade in the frame, namely the following range: from retention technology VTAS - 300 km/h to VNE = 350 km/h , till VTAS = 390 km/h. Flutter is non-existent in restricted range of calculated curve what is All before Ka-50 KAMOV's helicopters verified by flight test results. Calculated curve had full articulated rotor hubs. Ka-32 presents that there is a flutter speed margin of helicopter has metal hubs, mechanical & about 50 km/h (Fig.! 0 ). elastomeric bearings & dampers ( Fig.8 ). Ka-50 helicopter has metal and composite Vibration level of the coaxial helicopter hubs, elastomeric bearing & damper, flex have been discussed in paper [2]. The element for pitch, flapping and lag ( Fig.8 ). alternating forces apply to hubs of the upper/lower rotors and excite airframe 3.3 Rotor control linkage vibration. Configuration of the Ka-50 helicopter rotors affords applying of minhnal The rotor control linkage parameters alternating forces to the airframe. In this determine a flap-lag-pitch motion and a motion case vibration level of the airframe is stability of rotor blades ( Fig.6, Fig.8, Fig.9 ). minhnal too. The mathematical model of coaxial rotor control linkage was developed by KAMOV Vibration level is not in excess of O.Olg Company [!]. This math model is used for in main flight modes. Rotor pendulum and control linkage design and for frequency & anti-resonant isolation system are not used. stability analysis. The example shown on Fig. II [2]. The control linkage elasticity matrix Special task for the coaxial helicopter is functions was measured for full scale coaxial making a provision for the acceptable lower helicopters of four types. to-upper rotor blade tips clearances. As a task The analysis of the experhnental results of aeromechanics it is analogous with a task to determined to develop a control linkage provide the blade tips-to-tail boom clearance mathematical model and a adequate formula of the helicopters with the tail rotor.
G22- 3 KAMOV Company applies at analysis both 4.2 Ka-50 helicopter maneuverability calculation methods and flight tests [3, 8]. The features clearances are measured with the help of optic devices at each of 6 crossing points when the Load factor I speed envelope was upper blades are arranged above lower blades substantiated and validated by Ka-50 flight test during their relative rotation with doubled results: angular speed. - within operational limitations ( pitch, roll, rotor speed, rotor loads, ... ); The mechanics can be outlined as - within special aerobatic limitations. following ( Fig.l2 ). The planes of the upper I lower rotor blade tips are in parallel at A part of flight test points is illustrated by hover. Their clearance is even more than a the Fig.13: clearance between rotor hubs. at 3.5 > g-factor > 2 & at g-factor "' 0. At forward flight variable azimuthal airloads occur in the rotor disk, that results in Each point corresponds to one of the flapping motion. Because of this , planes of performed manoeuvres . The most part of the upper I lower rotor blade tips are inclined them are shown at the Fig.l3 No established to equal angles in flight direction ( forward I limitations have been exceeded . backward). Fig.l3 also shows the test flight results of Tiger's helicopter [13]. In lateral direction ( viewed along flight direction ) the planes of blade tips are The table on the Fig.l4 presents parameters inclined to each other because of counter- of manoeuvres within special limitations for rotation of the rotors ( Fig.l2 ). aerobatic flights. It is notable in this case parameters of "flat turn" and pull-out from The upper-to-lower tips clearances on one the skewed loop at g-factor = 3.5. disk side decreases and increases on the opposite one. In lateral direction an inclination angle of blade tip planes is 4.3 The means of aerobatic flight approximately equal to blade tip flapping angle monitoring and analysis ( to the left I to the right ) and depends on flight mode ( Fig.l2 ). As known from The NSTAR software was created to aeromechanics, there are relations between provide processing and analysis of Ka-50 blade flapping angle and the rotor parameters, helicopter test flight data. Using records made especially to Lock number, blade geometrical by aircraft test instrumentation the NSTAR twist angle and bladelcontrollinkage torsional software makes possible to restore the flying stiffuess. path and to calculate flight parameter Calculation and flight test results show the additional values [14]. values of coaxial rotor parameters mentioned NSTAR software is comparable both with above which ensure acceptable safety test flight record system and with standard clearance. record system. NSTAR results are used for Fig.l2 demonstrates measured blade tip the following pmposes: flapping angles made during flight tests of Ka- analysis of actions , assistance in pilot 50 helicopter and comparison with calculation training; data. examination for critical parameter limits; Ka-50 helicopter generalised measurement - as input data for simulation. results for the forward flight and manoeuvres Fig.l4 presents an example for skewed are presented in the Fig.l2, Fig.13. loop path recovery.
The acceptable upper-to-lower rotor blade 5. Conclusions tips clearances were substantiated by mathematical simulation and validated by 5.1 The basic aeromechanic & aeroelastic flight tests results for all approved envelope of problems of a coaxial rotor helicopter have manoeuvres. been developed and Key Technology have been achieved. The acceptable lower rotor blades to tail boom clearances were validated . 5.2 Key Technology determine the coaxial helicopter performance and manoeuvrability.
G22- 4 7. References 10. Eypuen, E.H., Baii:a:rmei!B, 11.M., KBoKoB, B .H.,Tie1pOCHH,3 A.," 7. Bourtsev, B.N., Gubarev, B.A., "A Ka- 15. Mikheyev,S.V.,Bourtsev,B.N.,Selemenev, 115 Helicopter a New Development of S.V., "Ka-50 Attack Helicopter Aerobatic KAMOV Company", Proceedings of22nd Flight", Proceedings of 24th European European Rotorcraft Forum, Russia, Rotorcraft Forum, France, Marseilles, Saint-Petersburg, 30 August - 1 Sept., 15-17Sept.l998. 1995. 16. Mikheyev,S.V.,Bourtsev,B.N.,Selemenev, 8. Bourtsev, B.N., Selemenev, S.V., "The S.V., "Ka-50 Attack Helicopter Aerobatic Flap Motion and the Upper Rotor Blades Flight", ASH 55th Annual Forum to Lower Rotor Blades Clearance for the Proceedings, Canada, Montreal, 25 - 27 Coaxial Helicopters", Journal of AHS, May 1999. No!, 1996. 9. Bourtsev, B.N., Kvokov, V.N., Vainstein, 17. CaMOXHH, B. ., EpMHJIOB, A.M., KoTIDip, I.M., Petrosian, B.A., "Phenomenon of a A.,[(., EypueB, E.H., CerreMeHes, C.B., Coaxial Helicopter High Figure of Merit "I1MrryJII:.CHOe aK)'CTWieCKOe H3.rrytieHHe at Hover", Proceedings of 23rd European BepTOJieTa COOCHOfi CXeM:DI IIpH Rotorcrafi Forum, Germany, Dresden, KpeHcepcKHX C:t G22- 5 Basic Parameters ( of Coaxial KAMOV'S Helicopters POWER LOADING Ww.x /P, [kg/h.p.] 8 7 MAX SPEED VMAX, [km/h] ( Fig.lA G22- 6 Basic Parameters of Coaxial KAMOV'S Helicopters RELATIVE HUB CLEARANCES DISK LOADING Fig.lB G22- 7 Single & Coaxial Rotors active disk areas , effective diameters , power & thrusts at hover SINGLE ROTOR COAXIAL ROTOR R R 0,85R 0,91R A A M=0,28A As=A=nD'/4 Ac =nD 2 /4+13A =1,28As DEFF = .j4Asjn = 1,13D Ts = (33,25 · 'ls · D · P.Ji )Y, Tc = (33,25 · 'lc · D · P.fi)Y, Single & Coaxial Rotor Helicopters main rotor diameter , power & thrust at hover HELICOPTER WITH A TAIL ROTOR COAXIAL HELICOPTER Ts Tc 0,9P 0,5P \ )"' 0,5P 7 p tG':\ -=r= I Pm-O,lP I \I rr I A p ~ I I I I ! Ls iLc I ' Ts = (33,25 · 'ls · Ds ·0,9P.Ji'{' Tc = (33,25 ·TJc ·De· P.Ji'{' From 'lciTJ,= 1,13 and Pc=Ps=P and PTR=0,1P: I. At Dc=Ds the thrust ratio is Tc/Ts= (1,13/0,9i13 =1,16; 2. At Ts = Tc the diameter ratio is Ds I De= 1,13 I 0,9 = 1,26. Fig.2 G22- 8 Figure of merit of coaxial rotors ( Flight test measured data ) llc 0.90 _j H o v e r '- Airframe Thrust & Power Loss I il I 0.85 1. was taken into account • ----- __ J______- , I ~--:---a----~~-~--~ -~ 1 0.75 - . ~' 1 ' 0.70 Ka - 32 advanced rotor 8 Ka -50 prototype II Ka -50 production -type A Ka - 32 production -type * MAX measured value 0.05 0.08 0.10 0.13 0.15 0.18 0.20 0.23 0.25 CT /cr Wake form in hover & its approximation r Ka - 32 COAXIAL ROTOR -_ .WAKE FORM APPROXIMATION: ,_ + -Upper Rotor : K1 ~ 0.0371 , K2 ~ 0.1028 j lx -Lower Rotor: K1 :;;:: 0.0431 , K2 = 0.0854 r I R = A + ( 1 -A ) I ch ( n A\If) ,.:f. Upper Rotor: A=0.82,n=2 A 2 tA 2 =Cr /CT =1.23 .. ";,{. 1 Lower Rotor: A= 0.91 n = 4 LR uR uR LR 0 30 60 90 120 150 180 210 240 270 300 330 360 WAKE AZIMUTH ANGLE , \jl deg Fig.3 G22- 9 THE STATISTICAL CHART ( Power Loadind - Disk Loading - Design Figure of Merit of Coaxial Helicopters & Helicopters with a Tail Rotor 2 2 Wmax I ( 1t R ) , [ kg 1 m ] 1 0 0 ------_-c-c--,-c:o------c- - - 0.5 0.6 ___o_._,7 ___ +_11_...:,c, ... 9 =_g._~ 8 ( Hover , A = 1 ) 7 2 Wmax) Wmax -- llc "'75. P nR2 ·A s-r--··~~-~~~~~-'~·~-1 "C 2 Wmax) Wmax c 5 1l s"' 75. 0,9·P nR2 ·A ·-"C Cll 0 ..J 4 ~ 1/) ·-c I ... "\ 0 0 -0:: c ·-Cll 2 ::!: 2 3 4 5 6 7 8 Wmax I P , [ kg I h.p.] - Power Loading Fig.4 G22- 10 Coaxial Rotor Wake Side Views for Several Flight Speeds / VrAs = 5 [km/h] t I I I I l VrAs = 73 [km/h] ... : .. ·· .. ·· VrAs = 138 [km/h] ... -·· VrAs = 227 [km/h] Wake Front Boundary Longitudinal Position Versus Flight Speed riR 1.0 -·- 0.5 2 3 3.5 4 5 r I R 0.5 . --=:::::::::::::::J::::::;;t;)-~#><~...... • v • I ... 1.0 _,.. I -=~~:j~=l~-~-~-~-~-2 1 - ·- 1 • I 1 1 ' { riR=1-21t!l, for riR > r0 IR, ~~3.5 r I R = r0 I R, for r I R ~ r 0 I R, V ~ 3.5 ,, ---·- r I R = 1 , for -V > 3.5 ll = VTAS·Cosal ( QR), V= VTAs 1..JT I ( 2pA) Fig.5 G22- 11 Simulated Aeroelastic Phenomena of Coaxial Rotor Simulated SIMULATION VERSION Phenomena ULISS-6 ULISS- ULMFE FLUT MFE Elx (r!R,mt) 1 EI, (r!R,mt) ./ ./ Glp (r!R,m t) - 2 Analysis Results of Coaxial Rotor Aeroelastic Simulation Analysis SIMULATION VERSION Results ULISS-6 ULISS-1 ULMFE FLUT MFE 1 Stall Coaxial Blade Blade flutter Rotors boundary 2 Bending moments, Coaxial Blade Blade Pitch link loads, Rotors Actuator loads 3 Elactic Coaxial Blade Blade Deformations Rotors 4 Alternate loads Coaxial on Hubs Rotors 5 Blade tips Coaxial Clearances Rotors 6 Flight Coaxial Blade Blade test flutter Rotors 7 Ground Coaxial Blade test flutter Rotors 8 Natural Blade Blade frequencies Fig.6 G22 -12 Blade Aerofoil Data Ka-50 Aerofoil 1·6 RAE9645 Anti Erosion Skin ~ 0 Trailing II Edge :2 1.4 ~ X Fibercarboglass Nomex ~ 1.2 +I ---!-'!."'-'.."'!~l8EiJi! Spar Honey Comb Heater Blanket Sweep 0.75 0.80 0.85 0.90 Woo AT (Ct_;Q) Blade Fitting Aft Section Aerodynamic moments of existing TsAGI-2 , cL TsAGI-4 airfoils & advanced TsAGI-4M airfoil - - TsAGI-2 airfoil 2.0 e e e e • - TsAGI - 4 airfoil ., .. 1 • " - TsAGI - 4M advanced ..... cw _.., -- Ka - 226 airfoil 1.5 ~ ~' 1.0 \ IJ 0.5 - : ' ' : 0.7·lJ. M=0.3 0.4 0.5 0.6 0.8 J ' 0 "'" )II Ill I' I I''' .. • • ...... 1 " o rnn n -0.2 -0.1 0.0 0.0 0.0 0.0 0.0 0.0 0.1 CM ( 0.25 Chord ) Fig.7 G22- 13 Ka-32 rotors control linkage model TOW TBAnf TnP TK man 1-Po -+-1- Po-+! ( Ka-50 rotors control linkage model TCt K2 TBAnt fTC Ai l..c;::=====!======::;:>.! Ai TnP TK Tnon f+-n ~I• n --1 I JO 1 JO I TOW Fig.8 G22- 14 The scheme of experiment on determination of the control linkage elasticity matrix ~lp. So= TOlll 2T0lll . (jl = !IJI,LR +Orr-~-~ (I -l); I=1,2, ... ,K 1 2TOlll 4TOlll + TBOlll ' 21t- (IJIILR +orr+ DFI)- ~-~(I- K-1); I= K+ 1, ... , 2K Main rigidities of the elasticity matrix & dynamic rigidities obtained from frequency testing 1 Mq5+S- -200 -100 0 100 200 '¥1LR, deg Fig.9 022- 15 Demonstrated Rotor Speed Range Blade Tip Mach Number ( JSA SLS) : 260~·~~~~~~~~·-~·--~~~~~0.9~8~~~~~~~--~--======-i 250 0 Q) ~ .§. 240 ~ "0 Q) Q) 230 0. (f) 0. i= ~ 220 0 0 -0:: 210 200 250 300 350 400 Stall Flutter Boundary CT/cr 0.5 "z" m.g CT/0' = 2 pt2·As·(.QR) '-"··-'.-f--· --o-· . Stall Limit ( Theory ) .... New Profiles DMH4, DMH3 ...... TIGER, EC135, 80105 KWS Cl s:: 0.3 -t-···-- i "'C Ka- 50 Flight Tests co Approximation 0 Ka -50 Flight Tests ~ 0.2 -t- ' <) TIGER, PAH2/ HAC co + TIGER,HAP al X EC135 Stress Flights -t-·- 0.1 0 WG13 Lynx Record !:;,. Dauphin DGV Test 0.0 0.1 0.2 0.3 0.4 0.5 0.6 0. ( Advance ratio, true airspeed !-! TAS [-] Fig.lO G22 -16 Ka - 25 Helicopter 3 CD Vertical Vibrations g, m/sek 2 0.25 BASIC 0.20 l IN THE CENTRE OF GRAVITY -··· - -··-·--·- ' . 0.15 :VIBRATION IREDUCTION 0 50 100 150 200 250 2 V, km/h g, m/sek 0.25 BASIC 0.20 0.15 0.10 VIBRATION REDUCTION 0.05 0.00 --+---.----,r--y---.,---...,---r---r--r---.-----; 0 50 100 150 200 250 V, km/h Fig.ll G22- 17 Blade Tip Coefficients Comparision of Calculations and Flight Test Results 500 400 .sE' o5 w1QQ -200 600 0 500 E' 400 .s 300 200 "' 100 0 600 500 E 4oo .s 300 0 50 100 150 200 250 300 350 VTAs [km/h] The Upper-to-Lower Rotor Blade Tips Clearancies Versus Level Forvard Speed & Blade Azimuth Flight Tests ! H,mm D 1500-2000 Approximation ! 2000 D 1000-1500 II 850-1000 1500 FLIGHT 47 !6 7 \If, deg VTAs, kmlh 250 300 Fig.12 022-18 Measured Upper-to-Lower Rotor Blade Tips Clearancies 1500 Ho 'E 1250 .s c ~ 1 QQQ MA~i~~~IES PERFORMED: ~ -HAMMERHEAD ~ -GLIDE C - DIVE & PULL-OUT ~ -HARDTURN gj 750 - COMBATTURN 13 - QUICK STOP "0 -ROLL-IN I ROLL-OUT ~ ~~-~s~~~~D~L~O~OP~~~B ~ 500 -= E z lilillll!l!l MEASURED CLEARANCIES FOR All ~ 250 i_-~--l------"--~ lilillll!l!l APPROVED ENVELOPE OF MANEUVRIES D THE RESERVED FIELD OF CLEARANCE I -150 -100 -50 0 50 100 150 200 250 300 350 400 'l-As [km/h] Load Factor I Speed Envelope ( Structural Qualification ) 3 ' ' ' ' o+----+-----+---I ,1 I -1 --1--r.--r-o--1~~~--~--~-,1 ~---l-r--,--r-o-l~-r-h~---l-r--,~h-r-,...,..-1-.---.-r-.--h--.---.-r-h-r-~ -150 -100 -50 0 50 100 150 200 250 300 350 400 VrAs [km/h] Fig.13 G22- 19 Ka-50 Aerobatic Maneuvries The Measured Parameter Values (minimax) ( MANEUVER Airspeed Pitch Roll DESCRIPTION VTAS Load Factor [g] attitude attitude [km/h] [deg] [deg] Hard Turn 280-;- 60 1.0 __... 2.9 __... 1.0 20-;- 50 0-;- -70 Unsteady Turn (Right/Left) with Pitch & Roll Flat Turn 220-;- 0 1.0 ---7 1.5 ---7 1.0 ±5 ±20 Jaw Attitude (Right/Left) ±80-.- ±90 [deg] Hammerhead 280-;- 0 1.0---7 2.9 ---7 1.0 0-;- ±90 ±90 (Right/Left) ---7 2.9 ---7 1.0 Dive 0-;- 390 1.0 ---7 0.25 0-;- -90 ±30 Push-Down, ---7 2.9 ---7 1.0 Dive & Pull-Out Skewed Loop 280-;- 70 1.0 ---7 2.9 ---7 1.2 0-;- 360 ±150 (Right/Left) ---7 3.5 ---7 1.0 Quick Stop 150-;- 40 1.0 ---7 2.0 ---7 1.2 0-;- 40 ±55 Pitch I Roll (Right/Left) ---7 1.0 Decceleration Pull-Up -90-;- 0 1.0 ---7 1.5 ---7 1.0 0 -;- -70 ±10 Backward Acceleration with the Tail & Pull-Up with the Tail Forward Forward/Up Flight Path While Performing Skewed Loop 14 ( \ Fig.14 t = 0 seK G22 -20