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, I NASA Technical Momorandum 84907
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1186-27290 (NASA-TtI-84907) HIGHLY liAlIEUVERABI.?; AIRCRAFI IECHNOIOGY (HIMA!) FIIGHI-FLUTTER !EST fRGGBAM (NASA) 20 p Be A02/MF AOl Unclas CSCL 01C 43299 G3/00 May 1984 .. ' ; ...
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National Aeronautics and \1ay 31,1£85 Review for gonoral rolaosa Space Administration NASA Technical Memorandum 64907 ,
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Michaol W. Kehoe NASA Amos Re:;sarch Center, Dryden Flight Research Facility, Edwards. California 93523 "
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Notional Aeronautics and Space Admlnlstr~lIon Am~3 R030arch Conter l. Dryden Flight Research Facility \, Edwards, California 93523 ; :'!,
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SUMHARY The highly maneuverable aircraft technology. (Hil1J\T) vehicle was evaluated in a - .
joint National Aeronautics and Space Administration and Air Force flight-test program.
The Hil1J\T vehicle is a remotely piloted research vehicle. Its design inco=porates the use of advanced composite materials in the wings, and canards for aeroelastic tailoring. A flight-flutter test program was conducted to clear a sufficient flight " envelope to allow for performance, stability and control, and loads testing. Testing was accomplished with and without flight control-surface dampers. Flutter clearance of the vehicle indicated satisfactory damping and damping trends for the structural modes of the HiHAT vehicle. The data presented in this report include frequency and damping plotted as a function of Mach number.
" SYMBOLS limit Mach number aeroservoelastic ASE primary control system PCS backup control system BCS
true velocity v control surface chord length c
angle of attack, deg a acceleration due to gravity g
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frequency, radjsec structural damping coefficient g Mach number M ItlTRODUt:TION The highly maneuverable aircraft technology (lli~~T) vehicle wa~ built by Rockwell International for use in a joint National Aeronautics and Space Admini~tration (NASA) and Air Force flight-test progrum. The vehicle was remotely piloted from a ground- based cockpit (ref. 1) and designed to demonstrate several advanced fighter tech- nologies, including: (1) close-coupled canards that provide aerodynamic benefits, (2) active controls using a digital fly-by-wire 87stem, (3) advanced composite and metallic structures, and (4) wing and canard aeroelastic tailoring.
The vehicle was launched from a B-52 aircraft at '13,716 m (45,000 ft) for each flight. The vehic:e weight at launch (full fuel) was 1587 kg (3500 lb). The overall vehicle dimensions are presented in figure 1. The cruise or maneuvel' camber of the wing and canard supercritical airfoils could be selected between flights by changing leading edges. The difference in the leading edges is illustrated in figure 2.
A flutter-clearance program was condu~ted to verify the freedom from flutter within the flight envelope for the HirmT vehicle. Clearance of the flutter envelope consisted of: flutter analyois (refs. 2 to 4), ground-vibration test (GVT) (ref. 5), wind-t~nnel test (ref. 5), and flight-flutter test.
Rockwell International, North ~~erican Aircraft Operations, performed the flutter- analysis and wind-tunnel tests. The wind-tunnel testing was performed only for the lJi ifSWhgr Ti£2t.~~~~-.-;.~ -:~:.l::':-:---- •.. 't·,·;,:,:::::,;~.",~",·! ..,..--.~ ~-r--"~'-r.~;'_r;' .. -'!'!.:--:;.:;._:;,;.= ....... - ... _..;-,.;;,-;.;.-;;;.-.;;;.--;;:_._;;;. .. _==...."..:::BII r--r~}'~ • I !
., investigaton of shock-induced oscillations. The GVT which was supported by NA~A and Rockwell International, was conducted by the Air Force Flight Test Center, Edwards AFB, California. Flight-flutter testing was conducted by the NASA Ames Dryden Flight Research Facility (DFRP) at Edwards, and supported by Rockwell International.
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Test Objectives The objectives of the program we~e to: (1) provide a flutter clearance for the HiMAT vehicle which would allow perfor~ance, stability and control, and loads testing to be accomplished, and (2) obt~in frequency and damping information for critical structucal modes of vibration.
Vehicle Description Two HiHAT vehicles were constructed. The first vehicle (Ship 1, serial no. 870) I .'
served as the basis for envelope expansion which included supersonic flight. The I ' second vehicle (Ship 2, serial no. 871) was used to probe the transonic region with
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, emphasis on pressure and structural deflection measurements. Ship 2 was not origi- nally intended to be flown supersonically during the flight-test program.
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, After the subsonic flutter clearance, Ship 1 was equipped with elevator and ele- I'" von control-surface dampers to prevent single-degree-of-frcedom control surface i .
flutter at supersonic speeds. These dampers were rc~oved toward the end of the ;. ~ .
flight-test program. Ship 2 was not equipped with control-surface dampers since the vehicle was not intended to be flown at suparsonic speeds. All flutter testing was l' conducted on Ship 1 (fig. 3).
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Two independent digital flight-control systems have been implecented in the Hi!~~T vehicles. Each vehicle was flo~m in the primary control system (peS) for normal research flight-test conditions. A backup control system (BCS) was provided to con- trol the vehicle i.n the event of certain airborne or ground failu~es which would pre- clude the use of the PCS (ref. 7). Transfer from the PCS to the BCS was designed to be automatic, but manual transfer could also be initiated.
The ailerons were mechanically locked out after the eighth flight, primarily as part of the new, relaxed, static stability-control. system testing within the flight envelope. It should be noted that the ailerons were predicted to flutter at super- sonic speeds, and may have required dampers had they not been locked out. The rud- ders, elevons, elevators, and canard flaps were active ~or all flutter flights
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(fig. 1). The rudders were all movable surfaces, while all the other cc.ltrol sur- .!\).
faces were hinged.
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Instrumentation Flutter instrumentation aboard Ship 1 consisted of the accelerometers and strain gages shown in figure 4. This instrumentation was used for all of the supersonic ana some of the subsonic flutter testing. Instrumentation for the initial subsonic test- ing did not consist of the right-hand elevan and elevator accelerometers. However, an accelerometer was mounted on the forward and aft tip areas of the left canard and wing for the initial subsonic testing.
TEST APPROACHES AND PROCEDURES Flight-Envelope Expansion· ..
Flutter testing was accomplished at altitudes of 12,192 m (40,000 ft) and 7,620 m (25,000 ftl. The maximum Mach numbers flown for data acquisiton were Mach 1.44 at 12,19~ m (40,000 ftl and ~~ch 1.29 at 7,620 m (25,000 ftl. Each flutter flight con- .
.. sisted of a constant-altitude, incremental Hach nur.ber flight-envelope expans:on •
The planned incremental test points are illustrated in figure 5. Angle-of-attack effects on damping were evaluated at selected transonic test conditions. The sub- sonic envelope was cleared, and then the supersonic envelope was cleared. Accom- plishment of the 12,192-m (40,000-ftl test points always preceded ac~omplishment of the 7,620-m (25,000-ftl test points.
Excitatioll The vehicle was excited at each test point by a sequence of control-surface pulses which were programmed into the ground-based computer. These pUlses provided symmetric and antisymmetric excitation using the ailerons, canard flaps, elevons, and rudders. The aileron pulses were deleted from the sequence for the flights when the ailerons were mechanically locked on the vp.hicle. The ~xecution ti,ne for the sequence of control-surface pulses was approximately 20 ncc. The d~plitude and duration of each control-surface pulse is illustrated in figure 6.
Random atmospheric turbulence was the primary source of excitation to the vehicle's structure. Typically, 1 min of data was collected at each subsonic test point, and 30 sec of data was acquired at each supersonic test point.
Envelope Expansion Procedure A consistent procedure was uaed during the testing of the HiHAT vehicle. T'ne pilot firat stabilized the vehicle at the test altitude and specified Hach nUmber~ When the vehicle was stabilized, the structure was excited by control-surface pUl~es and random atmospheric turbulence.
Telemetered data were displayed on strip charts and were digitally stored on a disk in the DFRF spectral-analysis facility. Selected accelerometer responses were monitored on a real-time spectrascope to provide frequency-dornain information. A :ourier analyz~r was used for auto-power spectrum calculations to obtain frequency and damping information from the random data. The data were smoothed by mUltiplying the aut,)correlation function by an exponential function. The half-power technique was usej to estimate damping. The logarithmic decrement technique was em~loyed to obtain frequency and damping information from the strip chart r~sponses to control-surface pulses. Clearance to the next higher Mach number test point was given by the test director in the spectral-analysis facility after the damping coefficients and trends for the structural modes were observed to be satisfactory.
Postflight analysis performed between flights consisted of calCUlating the auto- power spectrum for each accelerometer response to extract the structural frequency and damping values.
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TEST ~{ESULTS The aircraft-measured response to excitation was uned to calculate frequency and - .
damping values for each test point. As expected, the preprogrammed control-surface pulses could only excite the first nymmetric bending modes for the canard and the
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wing. Random atmospheric turbulence wao relied upon for excitation of the other higher-frequency structural modes. Howe~cr, there was often ir.sufficient atmoopheric '.
turbulence for good structural excitation because of the Hi!{AT high-altitude-only flight-test envelope. Therefore, it was difficult to conoistently excite ail struc- tural modes of interest with an adequate oignal-to-noise ratio. Analysis of modal data with a low signal-to-noise ratio reaults in scatter of the estimated damping values. However, random and control-surface pUlc& excitation were considered accept- able for this test program because (1) the flutter speed was predicted to be more than 20 parcant greater than the planned maximum speed test point, (2) speed was increased by a small increment between test points, and (3) the Hi~mT is a remotely piloted vehicle.
Subaonic data were acquired with both the cluioe and maneuver leading ~dges installed on the vehicle. All supernonic test pointB were accompliahed with the cruise leading edges only. The subsonic data results revealed no significant effects on the fluttor characteristica of the vehicle because of the leading-edge configura- tion chango.
The data also indicated that t~he installation of m9chanical locks on the ailer-
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ons did not significantly affect the critical modal characteristics of the vehicle. _.
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The effect of angle of attaclt on the modal dat1ping waa investigated. Turns at .'* ~,.
an elevated load factor and puohover-pul1up maneuvers uero performed in the tran- ~ ~~ sonic region to acquire the data. Analysis rave~lod that no definite trends could :~ be established because of scatter in the data.
The plots of frequency and damping as a function of Bach number contain averaged data. These data were obtained for e~ch mode by averaging ·the frequency and damping values calculated from several accelerometer responses. All frequency and damping data presented are from auto-po~er spectra analysea, these data are not presented for particular modes at some Mach numbers at which poor modal excitation was experienced.
Atmospheric turbulence sometimes failed to adequately excite a mode with a low energy content, resulting in a low reliability of ~le calcuiated damping value because the random excitation was of almost the same magnitUde as was the magnitUde of the noise.
The plots displayed were faired with least-squared error lines so that general data trends could be followed. The structural modes ~ere identified with names based on data in reference 5.
Test Data at 12,192 m (40,000 ft) The frequency 6nd damping trendo at 12,192 m (40,000 ft) are presented in fig- ures 7 to 12. No adverse damping trends were exhibited for the modes that were tracked. The maximum Mach number at which data were obtained was Bach 1.44.
The damping trend for the symmetric canard bending mode (fig. 9) exhibited a decrease in damping in the tranaonic region, however, the level of damping increaaed in the 3uperoonic region.
. , ~.i ~st oata at 7,620 m (25,000 ft) The frequency and damping trends at 7,620 m (25,000 ft) are presented in figures 13 to 18. The damping trends and lovels ,,-ere oatisfactory from a flutter standpoint.
The ml:4ximu:n Hach number at which dc3ta were obtained was Mach 1.29.
The flight-tent envelope presently cleared for the HHIAT vehicle is illustrated in figure 19. The envelope limits were establiched by the maximum Ibch number (1.44) and the maximum equivalent airspeed (518 KEAS) at which data were acquired. This envelope is less than the IUl1AT maximum design capability. More flight-flutter test- ing w~uld be required if further envelope exp~naion were desired.
Control Surface Free Play The first three flights of the vehicle were flown with excessive free play in the control surfaces. Flutter testing on the second and third flights revealed lightl~' damped oocillationa on the canard ~nd wing.
Extrapolation of tho damping trend for the canard indicated possible flutter at Bach 0.95 at 12,192 m (40,000 ft). Frequency and daI:lping trends of the wing and can- ard were obtained after flights 2 and 3 by autocorrelation/direct Fourier transform analyses performed by Rockwall International (ref. 8). Subsequent to flight 3, a flutter analysis (ref. B) with free play included was accomplished. The results of the analysis arc presented in' figure 20. Tho boundary correlates closely with the flight-test damping trends from flights 2 and 3.
The free play of all the control surfaces waD significantly reduced by modifying the control-llurface actuator bolts. Limits were sp3cificd (ref. 9) for each control surface (t~ble 1), and the free play was ce~Durcd after each flight to ensure that these limits were not exceeded. Flutter testing was conducted on flight 4, and a significant improvement in da~ping waa noted. Figure 21 shows a comparison of damp- ing trends and levels with and without free play. The damping estimates shown in figure 21 were obtained from control-surface pulses using the logarithmic decrc.;t1ent method (ref. 8). Note that the damping estimates exhibited in figures 13 (sYI:\metric wing bending) and 15 (sYI:\metric canard bending) are slightly different than the values shown in figure 21. The differences are likely due to the different tech- niques (logarithmic decrement and auto-powet' spectrum) of analysis.
\ Aerooervoelasticity During flight testing, the· Hi~~T vehicle was monitored for possible ap.roservo- elastic (ASE) instabilities. Such instabilities occur when the flight-control system dynamically interacts w~th structural modes. Adverse structural coupling of the flight-control aystem was observed with the first sYI:\mctric wing-bending mode (9 Hz) on two occasions. This coupling occurred in the pitch axis when BCS was engaged at Mach 0.88 at 11,278 rn (37,000 ft), and at ~mch 0.91 at 12,802 m (42,000 ft). r~ adverse coupling was noted while the vehicle was op~rated in the primary flight- control system.
Time history tr<:lces of the BCS coupling with the 9-Hz structural mode at Mach 0.88 at 11,278 m (37,000 ft) are presented in figure 22. Note that the wing oscillations are limited in amplitude. During the ncs engagement at ~~ch 0.91 at 12,802 m (42,000 ft), the vehicle was decelerated from f.l~ch 0.91 to Uach 0.80.
The 9-Hz oscillations remained constant in amplitude throughout tho deceleration.
Several other transfero to the BCS occurred at altitudes below 11,278 m (37,000 ft). No structural coupling was oboerved in the BCS. The maximum dynamic
proolJure at which the Des was engaged was 3.83 N/cm (800 psf) at Hach 1.2 and
7,620 m (25,000 ft).
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Control-Surface Dampers ... ~ Control-surf ace-rotation frequency measurements indicated that the control our- faces on the Hif.lAT vehicle had inoufficient restraint otiffness to meet Rockwell's single-degree-of-freedom flutter (buz%) criterion and could therefore be susceptible to buzz instabilities. This buzz criterion was established fro~ a Rockwell study and is summarized in figure 23. The criterion does not inClude tho effects of structural damping, angle of attack, and airfoil shape. The study indicated that buzz would not occur below a reduced velocity (V/cw) o~ 1.5, rcgardleso of P~ch nueber. The study also indicated that the mORt critical l-1ach nul'i\bar was approximately 1.25. Note that the unatable flight-test points are in the M~ch nur.~ar range of 1.0 to 1.1. It may
be inferred from thin study that the higher the reduced velocity io ~bove '.5, the
more likely that the control our face io to buzz.
Uoing this criterion, frcquency-verous-reduced-velocity curves were generated for the control surfaces on both HiUAT vohicloo (fig. 24). The rudders wero all- movable ourfacos. Tnerefora, the buzz criterion does not apply to tho ruddera on the HiMAT vehicle. When the measured control-aurfuce-rotation frequency values for both of the llLolAT ... ehicleo were plotted on figure 24, the I?.oc;nituda of tho frequency defi- cit was clearly indicated. An a result, detailed supersonic ccntrol-curf~ca flutter ...
analyses were required to oatoblioh rnech~nical dnmp~r rcquircwont~. The analyse3 aS6ulned no structural du.mping in the detcrfilin.:ttion of the control-surface viscoua dampers, pri~arily because of the unreliable nature of actuator system damping.
Thene analyses predicted that tha e~isting hinge-line rotational sti~fnogs was suf- ficient to prevent canard flap, single-degree-of-frecdom flutter, but that th~ ele- vons and elevators would roquire dampers.
Elevon and elevator damp~r8 were inotal1ed on Ship 1. The vehicle was flown to Mach 1.44 at 12,192 m (40,000 ft) with no indications of control-surface buzz at any Mach number.
It became desirable to fly Ship 2 supersonically so that wing and canard pressure distribution and deflection measurement data could be gathered. Ship 2 was not equipped with control-surface dampers. Becauae of money and time constraints, it was ducided not to modify Ship 2 with dampero, but to clear Ship 1 to fly supersonically with the control-surface dampars removed, thus clearing Ship 2 so that it could fly without control-ourface dampero. Ship 1 was flo~m without d~rnpers to Mach 1.2 at 12,192 m (40,000 ft) and to I}~ch 1.29 at 7,620 m (25,000 ft) with no indications of buzz. Ship 2 wac flown uithout dampers to Mach 1.2 at 12,192 m (40,000 ft), Mach 1.25 at 11,582 m (38,000 it) and Mach 1.2 at 7,620 m (25,000 ft) \Jith no indications of bu?z.
The buzz criterion and the reaults of the single-tegrec-of-freedom control-surface flutter analyses proved "to ~ conservative \"hen compared with the flight results of th~ Hi~~T with the dampers removed. The l~ck of agreement between the buzz criterion and the flight-test resulto may be a result of the inability of the criterion to '·x;;' .. nE'3-' • > account for the effects of structural damping, airfoil share, and angle of attack.
The lack of correlation b~twecn the analyneD and flight-test reoults may be due to the inaccuracies of the anolytical representation of the unGt~ady aerodynamics.
Another possible explanation io that the aosumption of no structural damping in the analysis was conservative. The rationale for this assumption is that as flight time ...
increases on a vehicle, the wear and tear on the control-system actuators can cause a significant reduction in the inherent structural damping. However, for the Hir'~T , I vehicle, the regular maintenance of the control system and the postflight free-play I checks would have prevented any significant reduction in. the actuator-system damping.
", CONCLUSIONS A flight-flutter test program wan successfully completed for the Hi~~T vehicle.
Data were acquired at Mach 1.44 and Mach 1.29 at 12,192 m (40,000 ft) and 7,620 m (25,000 ft), respectively. The damping levels and trends were satisfactory from a flutter standpoint. Three flights were flo~ with an excessive amount of free play I in the control surfaces. Lightly damped oscillations were recorded for the wing and I !
canard on the second and third flights. Extrapolation of the damping trends revealed a possibility of flutter at ~~ch 0.95 at 12,192 m (40,000 ft). SUbsequent'to the third flight, the free play was removed, and a significant improvement in damping was ~ ~ ~ exhibited on the fourth flight. J • ,~ \ .'
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HiMAT Ship 1 was flown to supersonic speeds with and without dampers on the ele- von and elevator control surfaces. Single-degree-of-frcedorn control-surf~ce flutter i was not exporienceu for either configuration, although the buz~ criterion and analy- ...
9C8 indicated that the damper-off configuration was susceptible to flutter.
i .~ OJ ., , A coupling of the flight-eontrol system with first sy~metric wing-bending mode ~i (9 Hz) was ey.perienccd at an altitude of 11,270 m (37,000 ft) and above while in the ~l BCS. The resulting oscillation was limited in amplitude. The BCS did not exhibit
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" this instability below 11,27B m (37,OOO·ft).
, i , ~ Ames Research Center ~ :: ,
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Dryden Flight Research Facll~t!J National Aeronautics and Space Administration
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Edwards, California, March 9, 1983 ..
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.. ' REFERENCES i i I' 1. Arnaiz, H. H., and Loschkc, P. C.: current OVerview of the Joint NASA/USAF HiMAT Program. NASA CP-2162, Part 1, 19B1, pp. 91-121.
- , 2. Siegel, S.: HiMAT Post-GVT Flutter Ana1yeis. NA-79-366, Rockwell International, North Americ~n Aircraft Div., Los ~~geles, Calif., Aug. 7, 1979.
3. Siegel, S.: Hil·mT Flutter Analysis. NA-78-708, Rockwell International, North American Aircraft Div., Los Angelos, Calif. Sept. 11, 1978.
4. Siegel,S.: Preliminary Flutter Analysis of the HiMAT/RPRV Air Vehicle.
DAR11420-01, Rockwell International, North 1merican Aircraft Div., '/ Los Ange1cs, Calif., Apr. 15, 1976.
5. Coombs, David F.: Ground Vibration Test of the Highly Man~uverab1e Aircraft Technology (lIiHAT) vehicle. AFFTC-TR-7B-31, Air Force Flight Test Center, Edwards, Calif., Mar. 1979.
6. Stevenson, J. R., Shock-Induced Self-Excited Airfoil Bending oacillations.
NA-78-723, Rockwell International, tbrth A~erican Aircraft Div., Los Angeles, Calif., OCt. 1978.
7. Hyers, Albert F., Earls, Itichac1 R., and ca11izo, Larry A.: HittAT Onboard I: • , .
Flight Computer System Architecture and Qualifications. AIM paper ..
~ 81-2107, OCt. 1981.
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A :2 8. Dobbs, S. K., and Siegel,S.: HiMAT Flutter CharacteristL::s with Control i ': .
Surface Frec Play: Flight Test and Ana1y3i3. TFO-80-J94, Rockwell International, North k~erican Aircraft Div., Los Angeles, Calif., Aug. 1, 1980.
; 9. Military Specification: Airplane Strength and Rigidity - Flutter, Divergence , .
i and Other Aeroelastic Inatabilities. HIL-A-008870~ (USAF), Mar. 31, 1971.
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OmGlNAL PAGE ro 0:: POOR QUALITY TABLE 1. - CONTROL SURFACE FREE-PLAY LIMITS Free-play limit, Control surface deg
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Aileron 0.13 Elevon 0.286 l.
Elevator 0.286 Rudder 0.034 Canard flap 0.13 Canard lIap Aileron Elevator \rUlse
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Figure 1. Hif!AT vehicle.
Dimensions in meters (feet).
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ECN 14281 Figure 3. UifMT test vehi=lp. in flight.
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Figure 6. Auto~a~ic flight ~ontrol pulse seq~bnce. Right control surface ~eflections were equal in mag~icude to the left surface deflections; the direction of the riJht surface der.lections ccrresponds to the sgrr::r.atry noted for ~a:.:1J pulse::.
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.8 .lI 1.0 1.1 1.2 1 r 1 1.4 1.5 MlIll:hnu'l'le.r M.ch numMf Figure 8. Antisy~~eLric win~-~naing Figure.: 7. Syr.:metric wing-bending
modal d~ta. Altitude = 12,192 m
modal data. Altitude - 12,192 m (40,000 ft).
(40,000 ft).
JI
ORIGINAL PI\G~ m
Or. POOR QUALIT'{ Le&c1!ng~ LUdlrl.~ed, o C .. II::.
o Cr .. l~o o McM\lftl' ., P.lln,uvlIr , .12
r
., Dlmpl/19,
D:mplno. .ee t-
:~:~
o
I,
o :t~Q~
I.
.0 ---II_.....----I_-I---'~--'--.l-..~ .
!
~ I!
I F~uoncy, 40 l::: Hz 20 ~o---o 0---00--0----0- I ; o I I I I ! I I I I I I I :-- .7 .a .ll 1.0 1.1 1.2 1.3 U 1.5 .II .0 1.0 1.1 1.2 1.3 1.4 1.5 ( Mach numll4r r:1Ch numbor i !- Figure 10. ,Antisymmetric canard- Figure 9. sycr"Qtric CltlUJrd-bending . .: modal d~t~. Altitude· 12,192 m ber.ding ll:Od<1l data. Altitude·
,"
----.. t:: (40,000 ft).
12,192 m (40,000 ft).
~ , ; I: " '-, ., Le«:Ir.(1 ~!13 \ ..
LGoC::nn~ .\ o eru;:o ...
o C;u!~ o r.~stlI;UYOf C rJMnr~ ~~ .~ .~
'12~
-:I " !.
:.
~
Dam:~ ::.0 _.....---L_..L----I_..l.._l.--1..--J
~ ..
I I I I I I I r.
a
1CO t~
' l:: "
," freqUGr.cy, 60 ~-o---o 1: ; i Froqumcy. uu Hz , '
.1
I Hz 60 I I I ! I I ~9101
t
.7 .a .9 1.0 1.1 1.2 1.3 U ,1.5 eo.
.a .9 1.0 1.1 1.2 1.3 1.4 1.5 Mach number Mcch numbltf
I·
Figure 12. r'linglct-bending codlll Figura 11. Canard tor:;ion 1!X>d~1 dntll.
i' data. Altitude a 12,192 m
Altitude • 12,192 m (40,000 ft).
(40,000 ft).
I
., I
I
!
" !; I, I' , , .
!
I
i
I.
/,._-- . 12 ORIG1Nr-,L Pt\Gi::: t~ OF POOR QUALnY IMldlng edge Lead!"'J odlJ3 o Crulo;e o CruIM • M;nauyet' e P,,!:mClUYef .12 - ..
°0
'12~
Damping, .~
Dcm~ng. .1l8 . c:o
~
o
o
.04 g .04 o Froqusncy. 40 b FrcqUCftCy. .. ...
Hz
Hz 20 ~ :f
~ I I I I ! I 0.
o .8 .; 1.0 1.1 1.2 1.3 7 .a ~ 1.0 1.1 Mai:h number Mach number Figure 13. !"/l:lIretric wing- F!gure 14. Antisymmet- benIing codal data. Alti- ric wing-banding modal tude ft 76fO m (25,000 ttl.
data. Altitude m 7620 m (25,000 ttl.
wdlngl'd::J L~ed!ng ;d<;)e o C:ulu o Cn:!~3 o Mf:MUYer ~ M~n:uvcr .12 .12 Dcmplng •• CD Dllmplr.g, .03 ii .04 .04 ~
F~nCY, :o~ ~I
I I ! I I I [. I I I I A ~ 1~ 1~ 1.2 1~ .7 .8 ~ 1.0 1.1 1.2 1.3 MIlCh nur..be!'
Mach number Figure 15. Syr:unetric caMrd- Figure 16. Antisym~etric bending modal d.:lta. Alti- canard-bending modal data.
tude - 7620 m (25,000 ttl.
Altitude - 7620 m (25,000 ft).
OIiUQJNAL PAG~ m
or POOR QUALITY
u:ldll'l'J odoo lced!"9~ o Cn;l:ll o Crul» o ManeuYlf .12 ~ 0 f;10MUYllr ",
o 0 '12~
Do1~: ~o----= Dlm:~,:o ~
_--III--.L.! _..J.'_...s...1 _.J.! _..J11-...J .0- I J I !--L-J
100 r
100 t
Fn:q~. ~
F~, COr
_--,'_.....l---J1l.---L--..L..-......L.' --II ~:eco-~ 60!.
Sl.7 .8 .9 1.0 1.1 1.2 1.3 1.4 .c .9 1.0 1.1 1.2 1.3 .7 MDcll "um!»r Mach numbGt' Figure 18. ffinglet-bending modal .'
Cal1.!1rd torsion Figure 17.
" .
diSta. Altitude· 7620 m (25,000 ft). t .>.::: Altitude a 7620 m t: __ J.
rxxliSl data.
l':!..:~ (25,000 Et).
~ ..
• !: " '.- .
lOx 10 &0 .ro A1l1tude.
II Allltud., 30 m ..
0'--.--JL...L. __ JL._-.l. __ ....I-_--J'-;-_-L-_---J 1.8 2.0 1.0 .6 .Il 1.0 1.2 1.4 Mach number Figure 19. Cleared flutter enveloF~'
. -
-, -~' , / ORIGINAL Pi\';;'~ ~ OF POOR QU.';Ln~{ lOx 10 .1 CO A1mudll, AJlltudlI.
It m ::0 II o 4 3 12 1i1 20 U~p1cd "~p rol!lt~n f~rocy, Hz Fisurc 20. Car.ard flutter-analysis results. Car.i::rd first banding and flap rotation; l! - 0.95; z;;atcb point solutions ~t g - o.
--0- "'11th Cl(~~'G fr~ J:tey (r.J.zMI 2 €:'::.I 3) --0- W~Jt C);o,.~'l'O froG pl.ciy (ii:o,'lt 4) .1!:O .125 .100
-0- _---.L1.---
D~~ • .075
.a-~~-~ a
g .oso .Q25 o L---L_-J.._ ........ _..L---'I----"--' a a a _0--0- D.m~ng. .trfS ~---o-- a g
8 8 a
.GSO .025 8 8@ 1 1 o L..-. • ..L _..1.---1.'_..1'_-'-_.1- ---II .T1S .600 li2!i .cro ms .COO .925 .G5O Mechnumw Figure 21. Sgmzr.etric wing- /100 canard-ber~ing modal d~ta compari- son. Altitude d 7620_~ (25,000 Eel.
'.
" ORlGIf"o'\L p,".~~ ';9 OF POOR QUALIY'f .. , , I ... :
/
..
, ; : ..
Ii
. "
.u -2 L..;0_.L.---L_..L---L_-'----I_-L.._L--I--1 o A Jl 1.2 1.0 z.o 2.4 :z.o 3.Z 3.0 4.0 1'1lM, co:: rJ.guro 22. 'rJ.c:J hJ.story of uJ.r.g ~c.tll~tJ.on. ECS ~l.tght control codel H • 0.00, ~tJ.tudo • 11,270 m (37,000 ft). .
"";;.;,"~i;f1i/;.2!k!:£~2iS::!fi~~S:~I:~}'1~'~.~;<·:~I~?":~ ~;t",;f;:';::~'+";7~?~.,O;'il:·~~:~~:~E~:~-";::-:~~~~~'v~ :::=f~~;~~~t:~~~"-';'~;;}7(.~;~'!!!'J~'i~;~ ~ ",.. ... r:;1 omG~NAL Ph\:'<o;;:. 1;Jl I I I OF POOR QUAU1"\'
a
FlullCl o ~U<:~ ftIecliy. 4 View
I
I' No lIutt"" I • I O~--"----"----~---'~---'----"'--""" I 1.1 1.0 D.9 ..
P1guro 23. Ccntrol-surface flutter criterion.
. -
F~~.,~ tlz : '.- Flgur~ 24. HiJ~1T buzz require~ent3.
N a 1.25/ a1titudo a 12,192 m (40,OOOft).
1. Ropott No. I 2. Government Acc" .. ion No. 3. RtClpitnt'l ClII1IC9 No.
NASA 'I'H-84907 4. RrpOrt On.
Tille and Sublltl. 5.
M&y (IUHAT) Highly Haneuverable Mrcraft 'l'achnolCXJY 6. PtflOlmi1\9 O't'lniution Codt !Ught-Flutter Teat Proqrall 7. Autl>o,lsl 8. "-forml1\9 Ortiniulion R.pott No.
H-l1e3 IUcha.l W. !(a hoe 10. Work Unit No.
9. p.,lorming Or~5nizalion Name and MQ ..
NASA Allee Reaearch Center 11. Conl'DCI 0' Orlnt No.
Drydon Flight Reaearch FaciUty P.O. Ilox 273 !:dwarda, Cditornia 13. Type 01 R.POrt and IVIed CovtrClc:l 12. SPonsoring ~y Name and Addrns Technical ~t'lorandua Space Mminhtra tion National Mlronautica and 14. SponIOl'1\9 At~c:y coo.
Waahington, D.C. 20546 nTOP 533-03-1 I 15. Supplomentary Not" -
-- ......... -
16. Abwact
I~
technoloqy Val The highly lIaneuverabl. aircraft I Hit'JlT) vehicl" eVl:luated in a joint N!.tional Aeronautics ond Spac. Ad .. ini.tratio~ Air Fo~co flight-teat prograD,
"""
The IUHAT vehicle is a re~otoly piloted recaarch vehicle. Ita cluign incorporl!tea the use ot advanced cO!ilpt1site ~atoriala in the win~., lind canllrda for aeroela~tic I: tallorin']. A flight-flu~ter tolt I'roqraa conducted to clear a aufficient fllght -a envelopa to allov for parfonaance, stability lind control, and loada teeting. Teatinq vaa acco=pUlhed with and vithout flight control-surface daClpGrs. Flutter clearance of the vehicle indicated catisf",ctory d"",pinq and trend. for the atructural da"pinq C10dea of the 1Ut'.AT vehicle. 'l1le daU preeented in thi. upon include frequency and "a .. ping plotted al a function of Hllch number.
"
•
17. Key \'lord. ISu9~ttd by Authorllil II:. Oi't'lbuti~ St.atement , ,..
I\e~otely piloted vehicle
Qii;";'}':";<:,j:;~
Flight flutter Control Gurface buzz
I
Control surfl:co fro'! play
Q2?i"-·~.G~~~
Aeroeervoelasticity PrlC' , 19. S(:'(U(It\" Cl.sssif. lof thiS ,.oort!
120. Secur,ty Clm,l. 101 this Pol,,' 2 No. of P.~.
1 1. In
Unelassified tJnclauitiGd Ie A02 . , ---~ .--~, .. ~,~, ' AJ.,:"'''''q,_ ._ ~~~.,-,- .~ .. .. '''',-. ... -.... ----.-:':;" .- I
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