Scale Inflatable-Winged Aircraft
N ASA/TM -2002-210721
Ground and Flight Evaulation of a Small Scale Inflatable-Winged Aircraft
James E. Murray, Joseph W Pahle, Stephen V. Thornton, Shannon Vogus NASA Dryden Flight Research Center Edwards, California
Tony Frackowiak Analytical Services & Materials, Inc. Hampton, Virginia
Joe Mello California Polytechnic University San Luis Obispo, California
Brook Norton Vertigo, Inc. Lake Elsinore, California
Jan uary 2002
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Ground and Flight Evaulation of a Small- Scale Inflatable-Winged Aircraft
James E. Murray , Joseph W Pahle, Stephen V. Thornton, Shannon Vogus NASA Dryden Flight Research Center Edwards, California
Tony Frackowiak Analytical Services & Materials, Inc. Hampton, Virginia
Jo e Mello California Polytechnic University San Luis Obispo, California
Brook Norton Vertigo, Inc. Lake Elsinore, California
National Aeronautics and Space Administration
Dryden Right Research Center Edwards, California 93523-0273
January 2002 ------
NOTICE Use of trade names or names of manufacturers in this document does not constitute an official endorsement of such products or manufacturers, either expressed or implied , by the National Aeronautics and Space Adminjstration.
A vailable from the following:
NASA Center for AeroSpace Information (CAS I) National TechnicaJ Information Service (NTIS) 7121 Standard Drive 5285 Port Royal Road Hanover, MD 21076-1320 Springfield, VA 22161 -2171 (301 ) 621 -0390 (703) 487-4650 GROUND AND FLIGHT EVALUATION OF A SMALL-SCALE INFLAT ABLE-WINGED AIRCRAFT
James E. Murray*, Joseph W. Pahle*, Stephen V. T horntont, Shannon Vogus:j: NASA Dryden Flight Research Center Edwards, Cali fornia
Tony Frackowiak§ Analytical Services & Materials, Inc. Hampton, Virginia
Joseph D. Mello, Ph~ California Pol ytechnic University San Luis Obispo, Cali fornia
Brook Norton** Vertigo, Inc. Lake Elsinore, Californi a
Abstract the inflatable wing. In-fl ight inflation of the wing was demonstrated in three flight operati ons, and mea ured A small -scale, instrumented research aircraft was flight data illustrated the dynamic characteristics flown to investigate the flight characteristics of during wing inflation and transition to controlled inflatable wings. Ground tests measured the static li fting flight. Wing infl ati on was rapid and the vehicle structural characteristics of the wing at different dynamics during inflation and transition were benign. inflation pressures, and these results compared The resulting angles of attack and of sideslip weFe favorably with analytical predictions. A research small , and the dynamic response was limited to roll quality instrumentation system was assembled, and heave moti o ns. largely from commercial off-the-shelf components, and installed in the aircraft. Initial flight operations Nomenclature were conducted with a conventional rigid wi ng having the same dimensions as the inflatable wing. a angle of attack, deg Subsequent flights were conducted with the inflatable angle of sideslip, deg wing. Research maneuvers were executed to identify normal accelerati on, g the trim, aerodynamic performance, and longitudinal center of gravity stability and control characteristics of the vehicle in CG its different wing configurations. For the angle-of longi tudinal stabil ity parameter, l/deg attack range spanned in this flight program, measured symmetric elevon control effectiveness flight data demonstrated that the ri gid wing was an parameter, l/deg effective simulator of the lift-generating capability of eN normal force coefficient normal-force curve slope parameter, l/deg * Aerospace Engineer, member. eNa t Branch Chief. COTS commercial off-the-shelf :I: Aerospace Engineer. § Aerospace Technician. EMI electromagneti c interference ~ Associate Professor. GPS Global Positioning System **Technologies Program Manager, member. g accelerati on of gravity Copyri ght © 200 I by the American Institute of Aeronauti cs and Astronautics, Inc. No copyri ght is asserted in the United States under mg vehicle weight, lb Title 17. U.S. Code. The U.S. Government has a royalty-free li cense to exerci se all ri ghts under the copyright claimed herein for NASA ational Aeronautics and Space Governmental purposes. All other rights are reserved by the copyri ght owner. Administrati on
American Institute of Aeronautics and Astronauti cs POPU pushover-pull up maturati on of miniaturi zed sensor technology, Global psig pounds per square inch gage Pos itioning System (GPS) receivers, and micro-controller hard ware by the electronics industry q dynamic pressure, Ib/ft2 has enabl ed resea rch-quality instrumentati on sys tems RJC radio control onboard small-scale vehi cles with onl y a modest reference area, ft2 weight, power, and cos t impac t.
Introduction This paper prese nts the res ults of ground and fli ght tes ts applied to a small-scale, research aircraft with an Inflatabl e structures have been considered for and inflatable wing. The inflatabl e wing and aircraft applied to a number of aeros pace applications. Earl y confi guration are briefl y described. Data from static des igners I considered press uri zed tubul ar structures to load tes ts are compared with analyti cal results for the carry some of th e aerodynamic fli ght loads. In the wing alone. Development of an inflati on system, wing 1950s, inflatabl e aircraft desi gns, including the 2 stowage and retention system, and research Goodyear Inflatopl ane -4 and the ML Aviation Utility5 instrumentation are described. Data from the onboard were fabricated using press urized airfoil shapes in research instrumentation system are used to compare which a noncylindrical shape was maintained by the trim, performance, and stability and control internal tensi on members. These low-pressure sys tems characteri stics of the vehicle when configured with the included ex ternal bracing to carry some of the inflatabl e wing and with a similar ri gid wing. Finally, aerodynamic loads. In the 1960s, a reentry vehicle flight data and ground-based photo images are used to concept6 was proposed usi ng inflatable tubul ar ill ustrate the dynamic characteri sti cs of the vehicle structures. Recent concepts include both baffled, during in-fli ght wing inflation and transition to segmented wing des igns7 and designs using multiple controlled lifting fli ght. Notice: Use of trade names or pres surized spars to roughl y define the airfoil shape names of manufac turers in this document does not and to carry the aerodynamic loads.8 M aterial and constitute an official endorsement of such products or fa brica tion advances have allowed current designs9. 10 manu facturers, either expressed or implied, by the to operate at hi gh inflation pres sure and support fully Nati onal A eronautics and Space Administration. cantilevered aerodynamic loads, and several applications have been demonstrated in fli ght. I I Inflatable Wing Description Inflatabl e wings produced for prev iously completed U. S. Navy research and development were made T he inflatabl e wings used in thi s program were availabl e to researchers at NASA Dryden Fli ght des igned and fabricated by Verti go, [nco (Lake Research Ce nter. These inflatabl e wings were Elsinore, Californi a) for a U. S. Navy program. The integrated into the des ign of two small-scale ( 15-25 inflatabl e wings fabricated for this U.S. Navy program Ib), instrumented, research aircraft confi gurations: a were prov ided to NASA Dryden at no cos t, and two pusher-powered conventional configurati on (1 -2000), research vehicles were des igned around th ese wings . and an unpowered winged lifting-body configuration. Fi gure 1 shows a si mplified schemati c of the wing. Onl y the results from the 1-2000 are contained in thi s paper. Conventional ground and fl ight test techniques The inflatabl e wing contains fi ve inflatabl e, were applied to this research aircraft to gain an cylindrical spars that run spanwi se from tip to tip. The understanding of the structural, aerodynamic, and spars are made of spi rally braided Vec tran threads (a operational ch aracteri sti cs of vehicles with Celanese AG product) laid ove r a ureth ane ga ba rrier. state-of-the-art inflatabl e wi ngs. A fabric webbing spar cap is aligned on the top and bottom of eac h of the spars. The wing span is 64 in. ti p Ground and fli ght tes tin g of inflatabl e structures at to tip, and the chord is 7.25 in. T he airfoil is a mall scale is attracti ve for everal reasons. M os t relati vely thi ck, sy mmetric sec tion NACA-002 1. The groun d and fli ght tes t operations are greatl y simplified wing does not contain any control surfaces. A mani fold when the mass of the tes t vehicle is low. Vehicle at th e ce nter of the wing hold the wing spars in fabri ca ti on costs, personn el co ts, and tes t range cos ts pos ition and prov ides a ri gid connec ti on between the are all red uced with smaller vehicles. Furthermore, the hi gh-press ure gas source ( 150 ps ig to 300 psig) and the
2 American Insti tute of Aeronautics and Astronauti cs .... _ --_. .. ----
Unidirectional spar cap-----... \ Rip-stop nylon Spirally braided outer skin Vectran~
Urethane gas barrier Open-cell foam High pressure spar '\ Manifold and w ing Wingtip \ attachment structure plate T ~~~\~~~iiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiiE 7.25 in. 1.- ~~~~~~ ~1 ·~------64 i n . ------;..~I
010459
Figure 1. Infl atabl e wing structu re. wing spars. Once in the manifold, the hi gh-pressure and wing-inflati on systems. The vehicle was configured gas passes into each spar through an inflati on pin that with a large, ri gid H-tail with large control surfaces (2 is mounted in the manifold. Between the spars and to elevons and 2 rudders) to enhance stability, damping, the trail ing edge of the wing is open-cell foam bonded and control authority, as well as to facilitate integrati on to the spars and to a rip-stop nylon outer skin. of the 1-2000 with a carrier aircraft for air-launched Additi onally, a rib at each tip ri gi dl y connects all the operati ons. Because the inflatabl e wings had no control spars to establ ish wing torsional stiffness. Therm ally surfaces, full three-ax is control was effected onl y by the ac tivated ad hes ives are used to bond the spars, foam, tail control surfaces; the symmetric elevon controlled and the nylon skin into a contiguous wing structure. pitch, the differenti al elevon controlled roll, and the symmetric rudder controlled yaw. I-2000 Ve hicle Description
T he 1-2000 was capabl e of fli ght in anyone of three To evaluate a small-scale inflatable wing, a research wing confi gurati ons: rigid wing. a conventional aircraft des ignated the 1-2000 was designed and built. foam-and-fiberglass wing using geometry identical to T he 1-2000 research vehicle, shown in fi gure 2, is a that of the inflatabl e wing, preinflated wing. a wing fa irly conventional aircraft configurati on. This vehicle inflated on the ground prior to fl ight, or in-flight was designed to max imize operational flex ibility and the inflated wing. a wing capable of inflation while in quality of resea rch data obtained in the fl ight program. fl ight. Conversion among the three w ing The vehi cle was designed for fl ight as either a powered configura tion capable of conventional takeoff and configurations was fac ilitated by fabri cating multi pie landi ng, or as an unpowered glider configuration wing-deck assemblies to mate with the fuselage capable of being air-launched from a separate, powered assembly. The fuselage assembly contained the carrier aircraft. The powered 1-2000 was designed as a primary aircraft systems, while each wing- deck pusher to leave the nose clear for an airdata probe and assembly held the remaini ng systems req uired to maxi mize th e quality of the airdata measurements. The support the specific wing co nfiguration (e.g. inflati on fuselage was made large and boxlike to allow the system hard ware). Longitudinal center-of-grav ity freedom to install th e onboard systems, incl uding (CG) locati ons were identi cal for all configurati ons, ins trumentation , fuel tanks , uplink control hard ware, although the vertical CG locati on did vary wi th
3 American Institute of Aeronautics and Astronautics configuration. Vehicle weight ranged from 11.0 to 15.7 Ib th ro ughout the fli ght program.
Inflatable Wing Structural Testing
To structurally characteri ze the inflatabl e wing in preparation for fli ght tes ting, a series of static load tes ts was conducted. The wing was mounted at the centerline by cl amping the inflation manifold in a ri gid fi xture (fi g 3). Wing inflation press ure was supplied by regulated gaseous nitrogen. The loads were applied symmetrically and vertically at the wingtips using linear electromechanical actuators. Preliminary tests were conducted to determine the shear center of the wing. The actuators were then moved to the shear center position at the wingtips to induce a bending load with no torsional component. The applied loads were measured using load cells and recorded on a personal computer-based data acquisition sys tem. Wingtip defl ections were monitored with linear displ acement sensors.
Elevon control surfaces
64.00 in. 6.50 in.
Fi gure 3. Stati c structural testing of th e inflatabl e wi ng.
Loading tests were conducted usi ng three different Rudder wing inflation pre ures : 150 psig, 225 psig, and 300 control surfaces psig. Fi gure 4 presents the te t results for th e left wing panel. 8 eginni ng at zero load and zero defl ection, there is a characteristic and almost linear increase of load wi th increasing defl ecti on for th e first portion of the _~_ _; __~701;; curve, followed by a significant reducti on in slope out to the max imal load and defl ec ti on. T he return path to th e unloaded condition creates a hys teres is loop, with load I. 51.38 in. ----.~I 01046 being somew hat less for the decreas in g load condi tion than for the increasing load cond iti on at the same Fi gure 2. 1-2000 research vehi cle. deflection. The physical mechani sm that crea tes th e hys teresis loop is unknown. Visual inspection during
4 American Institute of Aeronautics and Astronautics researchers 12-14 have successfully employed mechanics 30r------~ I I I _ of materi als methods to inflated structures similar to the 25 -----? inflatabl e wing; th is work was limited to single tubes or ..,. ... -- ..··· ····-·1 structures in general. In the present work these methods 20 I------l-----I--:cI ...... --- •.. , _ .•. - I were extended to the mUlti-s par confi gurati on of the ..... ~.J . ,..,. .... _ _ - Load' 15 i ' .~- inflatabl e wing. Also, previous work employed Ib ,. "".~ - 1-/ •• homogeneous, isotropic, constant cross section •••1/ I '" 10 /7 -~;.---t---i Wing inflation structures. Becau e the compos ite micromechani cs of ;. 'b pressure I the material and structure of the inflatabl e wing spar was 150 psig _ 5 225 psig more complicated, and because of the complex and • 300 psig progress ive nature of the 5-tube response, the governing o 2 3 4 5 6 7 8 9 equati ons were coded in a MATLAB script fi le to Deflection, in. 010462 analyti cally predict the behavior. Fi gure 4. Left wingtip load as a function of wingtip defl ecti on. Figure 5 shows these results from the mechani cs of materials analytical approach compared to test data. The the testing confirmed that wrinkles in the spar tubes model captures three salient characteristics of the test formed (or relaxed) at the wi ng root during the peri od of data: the pre-wrinkle or initial linear slope (which is slope change. independent of inflati on pressure), the slope change at onset of wrinkling, and the linear increase in intial Inflatable Wing Structural Analys is wrinkle load with inflation pressure. The model slightl y underpredicts the stiffness of the structure in th e linear A brief analytical study was conducted to region and overpredicts the pos t-wrinkle defl ecti on. complement the inflatabl e wing tes ting. The purpose of From these res ults, it appears that the inflated wing thi s work was to investi gate analytical and computati onal structural models that mi ght be structure can be modeled effecti vely. A mechanics of applicabl e to this type of structure. The res ults offer materials type approach seems robust and is some insight into wing behav ior and some appropriate recommended for preliminary wing des ign. T he analysis and modeling techniques. methods developed here could poss ibly be extended to
During the structural testing, inspecti on of the Test Analytical structure while under load and a study of tes t data led to data results ~ 150 psig o 150 psig the following observati ons. These are key in '* Onset of ¢ 225 psig x 225 psig wrinkling understand ing the behavior and the subsequent o 300 psig • 300 psig development of appropriate modeling techniques. 25 ,------,
• The initial (linear) stiffness of th e wing is nearl y the 20 1------..<1'1"'" same throughout the range of inflation press ures tested. 15 1---- • T he load at whi ch the onset of wrinkling occurs Load, appears to be a linear function of inflati on pressure. Ib 10 • Inspection under the wing covering and foam in the root reg ion, during tests, revealed that the spar caps in the upper tubes wrinkle progressively in the 5 nonlinear ra nge of the wingtip load as a functi on of displacement. o 2 4 6 8 10 Two basic modeling approaches were inves ti gated: a Deflection, in. 010463 mechani cs of materi als analytical approach and a fi nite element approach. Onl y the res ults of the mechanics of Fi gure 5. Analyti cal res ults compared to tes t data for materials analytical approach are presented. Many win gtip load as a function of defl ection.
5 American In stitute of Aeronautics and Astronautics include torsion and the superpoSItIon of bending and vessel was mated with a COT S adjustable regul ator that torsion. However, it may be more economical to included an integrated fi ll port , pressure relief plug, and investi gate simplified finite-element models for these manually-actuated source valve (fig 7). The output of other loading types. th e integrated regul ator assembl y was connec ted to the wing inflation mani fold. Dry nitrogen was used for all To accurately model the nonlinear res ponse of uch a ground and fl ight tests. structure beyond the onset of wrinkling would require additional testing and computational development. A 35in3 ./" special ized finite-element model and a materi al model ./ pressure ./"./" may be required to deal wi th the inherent numerical / vessel stability problems for such a structure. ~
Inflation Gas Subsystem Design and Testing
Selecti on of wing inflation press ure was based on the res ults of the static structural characteri zati on of the inflatabl e wing. The wingtip load corres ponding to the onset of wrinkling was determined for each inflation pres sure tested (fi g 5). A ss uming an elliptical wing lift distribution and a 15- lb vehi cle gross weight, the vehicle load factor corresponding to the wingtip load at the Fi gure 7. Inflation gas subsys tem press ure vessel and onset of wrinkling was calcul ated. Figure 6 shows the regulator. vehicle load factor at onset of wrinkling as a functi on of inflati on press ure. Based on th c c res ults, a minimum T he same inflati on gas subsystem was used for both wing inflati on pressure of 180 psig was selected for pre- inflated fli ght and for in-fli ght inflation operations. mos t fli ght operati on to allow for a 3.5-g envelope. When confi gured for pre- inflated fli ghts the wing was slow ly inflated on the ground and onl y th e fi nal wing 7 press ure was important. For these flights, the reg ul ator 6 press ure was set at the des ired wing pressure (180 to 5 240 psi g), and the hi gh-press ure source tank was Load pressurized before fli ght to approximately 500 psig. factor 4 at T he exces gas in the hi gh-press ure source tank was onset of wrinkling, 3 th en availabl e during fli ght to make up any losses in the g 2 sys tem res ulting from leakage.
When configured for in-fli ght inflation, the inflati on o gas subsytem was required to control both th e fi nal wing 150 200 250 300 350 Inflation pressure, psig pressure and th e wing inflati on rate. In this 010464 confi gurati on, the adjustable regul ator was effec ti vely Fi gure 6. Allowabl e load factor as a f uncti on of wing used as an adj u table ori fice and the wing inflation inflati on pressure. sys tem was a blow down (unregul ated) sy tern . Final wing pressure was controlled exclu ively by the ini ti al Laboratory testing meas ured th e wing leak-rate under press ure in th e hi gh-press ure source tank; an initial tank the expec ted fli ght load, vibrati on, and temperature pressure of approximately 1800 psig would yield the conditions. T he results allowed appropri ate sizing of the desired final wing (and tank) press ure of approximately onboard inflation gas subsystem for the ex pec ted fli ght 180 psig. Mass fl ow rate, and thus wing inflat ion rate, durat ion. A small comm ercial off-the-shelf (COTS) was strongly dependent on th e regulator set point, and press ure vessel with a volume of approximately 35 in3 th erefore wing inflati on rate was controllabl e by means was selected as th e hi gh-press ure source tank. Thi of the regul ator press ure set poi nt.
6 A merican Institute of Aeronautics and Astronau tics -----~- .~--
Laboratory testing was used to find the regulator set point corresponding to the desired wing inflation rate. In order to limit the number of inflation cycles conducted with the actual wings, a rigid pressure vessel with volume equivalent to the inflatable wings was used as a wing simulator. Figure 8 shows the pressure time history within this wing simulator as a fu nction of the regulator pressure set point. The maximum allowable inflation rate was specified by the wing manufacturer. The desired inflation rate was determined from simulation, indicating the required load factor as a Figure 9. 1-2000 re earch vehicle mated with air-launch function of time for a pullout from a ballistic trajectory. carri er aircraft. Based on these test results, a regulator set point of 500
psig was selected for the in-flight inflation operations. wings in both stowed and inflated configu ration. Each wing panel was z-folded from the wingtip and the Maximum allowable Minimum allowable stowed structure was retained along the side of the inflation rate inflation rate fuselage with a horizontal fabric strap. Each fabric strap 200 180 160 140 Wing 120 - Regulator pressure, 100 set point, psig psig 80 200 60 300 40 400 cmnmo 500 20 600 o 2 3 4 5 6 7 8 Time, sec 0 10466 Figure 8. Wing inflation pressure time hi story as a function of regulator set poi nt.
Wing Stowage and Retention Subsystem Design
For in-flight inflation operations, the 1-2000 research vehicle with its wings deflated and stowed was carried to its release altitude mated with the air-launch carri er aircraft (fig 9). A system was required for stowing and retaining the deflated wings while the research vehicle was mated with the air-launch carrier aircraft and while
the research vehicle was in ballistic flight prior to wing 010467 inflation. For the 1-2000, there was no requirement for the deflated wings to be stowed wi thin th e body of th e Figure 10. Photo comparison of 1-2000 with wings veh icle. Figure 10 shows the 1-2000 vehicle with the stowed (top) and inflated (bottom).
7 American Institute of Aeronautics and Astronautics
------~~~~ was fixed to the fuselage at its front and was terminated Extensive ground testing was used to adjust the wi th a loop at the aft end. Each loop end was retai ned relative timing, using the wing simulator to replace the by a pin driven by a small pneumatic actuator mounted actual wing test article. Finally, a single ground test of on the fuselage just aft of the stowed wi ng assembly. the integrated in-flight inflation system was done for fI ight qual i fication. Infl ation System Integration and Testi ng Ai rborne Systems and Instrumentation
The inflation gas subsystem and the wing stowage The research vehicle was equipped with a COTS subsystem were integrated to form the complete wing command-uplink radio control (RfC) system. The inflation system. A schematic of the integrated system ground research pilot kept the research vehicle in direct is shown in fig ure II. The primary objective of the sight throughout each flight operation, and controlled all wing inflation system integration was to rel iably control aspects of the research mission with a COTS uplink the relative timing of the wing-retention pi n release and control computer-transmitter. Control surface gains, the wing inflation valve opening. The timing objective throws, and interconnects, as well as stick shaping and was for pin release to occur 100 msec (±50 msec) prior trim capability were available to the research pilot through the computer-transmitter. Onboard systems to valve opening. Two small pneumatic cylinders included a receiver-computer, conventional RfC actuated the wing retention pins and one larger servoactuators, and redundant battery power systems. pneumatic cylinder actuated the wi ng inflation valve. No additional stability augmentation or rate damping Two small mechanically driven spool valves controlled was implemented onboard the research vehicle. the flow of low-pressure (120 psig) actuation gas to the pneumatic cylinders. Two small servoactuators drove The vehicle was instrumented for flight dynamics, the spool valves. Relative timing of wing-retention pi n performance, and subsystem health measurements. The release and wing inflation valve opening was controlled core of the instrumentation system was a small COTS by modifying the relative timing of the command signal single-board microcontroller-based data-logging to the separate servoactuators. engine. This system wa supplemented with power
High pressure N2 tank (35 in3) (1800 psig) High pressure O------1~------_; fill port Wing L..-_--' inflation pneumatic cylinder Fill ports o--t-D-+----J Wing release servoactuator Wing release Wing inflation spool valve servoactuator
Left wing Right wing retention retention pneumatic pressure pneumatic cylinder regulator cylinder r.:-----, Wing manifold ....-----1 pressure transducer
010468
Figure 11. Schematic of the integrated in-flight-inflation system.
8 American Institute of Aeronautics and ASlronautics
i _J conditioning and signal conditioning circuit board s suspension geometry. Three separate orth ogonal appropri ate for the analog transducers used. During suspension orientations were used to identify the each fli ght operation, resea rch data were logged to important components of th e inerti a tensor. T he onboard sys tem memory, and the data were downloaded different suspension orientati ons allowed compari son of to a laptop computer at the end of each flight for further the meas ured inertia tensor components from different processing and analysis. T here was no downlink system experiments, improving confi dence in the results. for the flight data.
Instrumentation selection was driven by availability and the desire to minimize weight, power required, and cost. All instrumentation components were COTS units. Each control surface position was instrumented with a co ntrol-pos iti ~ n transducer. All airdata meas urements were made with a small airdata probe. On the probe, angle of attack (a ) and angle of sideslip ( 13 ) were instrumented with vane-driven potentiometers. Pitot and static ports on the probe were plumbed with tubing to absolute (static) and differential (pi tot minus static) piezoresisti ve press ure transducers mounted in the vehicle body. Body-axis angular rates were measured with ceramic Coriolis-effect rate tra nsducers, and body-axis accelerati on measurements were made with a triaxial piezoresisti ve accelerometer package. Vehicle attitude was not direc tl y meas ured; pos tflight traj ectory reconstruction was used to synthesize vehicle pitch attitude during wings-level flight by using measured altitude rate and a. High-pressure tank and wing inflation pressure measurements were made with pi ezores isti ve gage pressure transducers fabricated in Fi gure 12. Test configuration for inertia sw ing on stainless steel enclos ures. 1-2000 (suspension lines exaggerated for cl arity).
Prior to th e initiation of fli ght operations, the The onboard data system was used to record electromagnetic interference (EM!) susceptibility of the body-axis rates durin g the suspension ex periments, and upl ink-command system to the additional onboard parameter estimation techniques were used to es timate systems was meas ured through a standard range- test the inertia components. All inertia components were procedure. Initial range tes ting identified the need for corrected for the mass of the suspen sion hardware used. an EMI-s hi elded enclosure on the instrumentation Tabl e I presents the inertia swing results for two of the system, which was then implemented. fli ght configurati ons.
Vehicle Inertia Swings Tabl e I. M eas ured research vehicle inerti as .
Analysis of fli ght data and development of a Inert ia Rigid Wi ng Configurat ion simulation req uired accurate meas urement of the Component Average, Standard Deviati on, vehicle inertial properti es. A bi fi lar pendulum sl ug-ft2 slug-ft2 l 5 suspension technique was used to ex perimentally Ix 0. 1701 0.0 11 2 measure the vehi cle moments of inertia and the cross Iy 0.7647 0.0046 prod ucts of inertia. The bi fi lar suspension approach Iz 0.8776 0.0018 (fig 12) allows four degrees of freedom and allows -0.0009 0.0038 sim ultaneous identi fication of multiple moments of Ixy inertia and cross products of inertia with a single Ixz 0.05 15 0.0086
9 American Institute of Ae ronautics and Astronautics Table I . continued. series, the vehicle weight was also increased in one pound increments until the maximum expected flight Inertia Preinflated Wing Component Configuration weight was reached. Finally, the powered, ri gid-wing configuration was used to simulate and practice the Average, Standard Deviation, slug-ft2 sl ug-ft2 maneuver sequence planned for the air-launched, unpowered , in-flight-inflated configuration flights. Ix 0.2290 0.0101 Iy 0.7186 0.0010 Following the initial series of research flights in the Iz 0.8970 0.0045 powered, rigid-wing configuration, the vehicle was Ixy 0.0045 0.0062 modified and flown in the powered, preinflated-wing Ixz 0.0378 0.0106 configuration. For these research flights, the inflatable wing was inflated on the ground several minutes prior to takeoff, and the onboard pre sure systems were used Flight Test Approach to maintain wing pressure at approximately 180 psig. All flight piloting was performed by an expert-class During this preinflated flight series, research ground-based pilot using line-of-s ight visual cues only. maneuvers included longitudinal doublets for stability Ground and flight operations involving pressurized and control derivative estimation, and POPUs for trim systems were briefed beforehand, and safety zones and performance measurement. around the aircraft were restricted to essential personnel. After all research objectives were met with the 1-2000 in the powered, preinflated-wing configuration, The flight test approach of the 1-2000 followed a the vehicle wa prepared for unpowered air-launched conservative build-up approach commonly used in flights with in-flight inflation of the wing. The engine developmental flight testing. The objectives of the and all associated hardware were rem oved from the i nitial flights were to develop the operational vehicle and the in-flight inflation sys tem hardware was procedures, to wring out the airframe (adjust control installed. A hook was installed in the top of the system gains and control surface throws, optimize wing-deck assembly for mating to the belly of the engine performance, adjust gear geometry), and to air-launch carrier aircraft. One captive-carry flight check out the instrumentation system. The initial was conducted in the mated configuration to practice flights were made in the powered configuration and at ai r-launch operational procedures and to confirm the minimum weight in order to minimize flight loads and release-poi nt flight conditions. Following the to minimize takeoff and landing speeds. Furthermore, captive-carry flight, three flights were made with air for the initial flights the vehicle was configured with launch and in-flight inflation of the wing. Becau se the the rigid wing in order to eliminate risk to the unique duration of the5e un powered flights was short, no i nflatable flight-test article. intentional research maneuvers were performed.
Following the initial checkout flights, several Flight Data Results research flights were flown in the ri gid-wing configuration. The objective of these fli ghts was to Rigid Wing Compared to Inflatable Wing document the trim, performance, and stability and control characteristics of the 1-2000 ve hicle in its The avai lable flight data allowed comparison of the baseline rigid-wing configuration. These research lift-generating capability and the trim curve of the flights were also used to develop and practice so me of ai rcraft in three different configurations: th e ri gid the flight-test maneuvers planned for flight with the wing, the preinflated wing, and the in-flight inflated inflatable wing. During these research fli ghts, the wing. The fir t two configurations were powered while in trumentation system collected data for postflight the last was unpowered. To minimize the effect of analysis. The research maneuvers executed included unknown (i .e. unmeasured) engine thrust, normal force doublets for stab ility and control derivative estimation, coefficient was used for the comparison rather than lift and pushover-pullup (POPU) maneuvers for trim and coefficient because the thrust ax is was perpendicular to performance measurement. During this first fli ght the vehicle normal axis.
10 American Inslilule of Aeronautics and ASlronaulics To make the comparison of the different the flight data show that the lift-generating capability of configurations most meaningful, only selected subsets the vehicle is repeatable across the three configurations. of the flight data were used. There were three criteria These data also demonstrate that the rigid-wi ng for selecting the flight data for this comparison: power configuration can be an effective simulator of the setting, symmetric elevon rate, and roll rate. For the inflated configurations. powered configurations, only flight data with idle-power throttle settings were used; for the Figure 14 shows a comparison of the vehicle trim unpowered configuration, all data were available. For curve (a as a function of sy mmetric elevon) for the flights with POPU maneuvers flown, that portion of same flight data subsets shown in figure 13. For the each POPU with a smooth and slow (target of 1 deg/sec) symmetric elevon rate during the pull up Symbol Flight Configuration portion were used. Similar criteria were used to screen number flights that did not contain intentional POPU 0 26 Rigid wing x 27 Rigid w ing maneuvers. Finally, only flight data with small roll ~ 30 Preinflated wing rates were used. Given these selection criteria, 0 34 In-flight inflated portions of four flight data sets were available for wing 10 comparison.
8 Normal force coefficient was calculated from the flight-measured normal accelerometer and dynamic 6 press ure measurements, and the configuration-specific ex, vehicle weight: deg 4
( I ) 2
0 . - ' Figure 13 shows a comparison of the vehicle normal-force coefficient as a function of a for the three -2 configurations. For the a range spanned in the analysis, -3 -2 - 1 0 2 3 4 5 Symmetric elevon, deg 010471 Symbol Flight Configuration number Figure 14. Trim curve for three confi gurations. o 26 Rigid wing x 27 Rigid wing trim range spanned in the analysi s, the rigid-wing and ~ 30 Preinflated wing o 34 In-f light inflated preinflated-wing configurations show steeper (larger wing .9 negative slope) trim curves than the in-flight-inflated configuration. The steeper trim curves are consistent .8 with larger symmetric elevon effectiveness. Given that .7 '" figure 13 showed no significant difference in the .6 lift-generating capabi lity of the three configurations, .5 th e source of the trim curve difference is more likely CN .4 attri buted to the removal of the engi ne rather than a .3 difference in the aerodynamics of the in-flight inflated .2 wing. By removing the engine for the in- fli ght-inflated configurations, the entrained fl ow over the control .1 surfaces was reduced , thereby reducing their o 2 4 6 8 10 effectiveness. ex, deg 010470 Roll trim of the rigid-wing and preinflated-wing Figure 13. Normal force coefficient as a function of a configurations was measurably different. With respect for three configurations. to the ri gid-wing configuration, the initial flight with the
II American Institute of Aeronautics and Astronautics preinflated wing required approximately an additional Figure 15 compares the flight-measured and 10 degrees of differential elevon to trim. Postflight computed time history for one longitudinal doublet measurements of the wing revealed that when inflated, maneuver analyzed for the rigid-wing configuration. the wing had a small amount of twist unintentionally Owing to the large horizontal tail and the low flight built into each wing panel. For all subsequent flight speed, the short-period mode is heavily damped -- no activity, a small trim tab was affixed to the left wing free oscillation is apparent after the pilot control mOli on panel to correct the roll trim. (symmetric elevon) is stopped. This maneuver, performed in level flight with significant engine thrust Parameter Estimation Results for Simulation Development maintaining speed and altitude, provided good estimates of the primary stability and control parameters- The initial aerodynamic model used for simulation normal-force curve slope parameter (C N ), a was developed analytically using a vortex-lattice panel ), longitudinal stability parameter (Cm and symmetric code 16 . An update of this initial aerodynamic model elevon control effectiveness parameter ( C m . ) . 6. with flight-derived results was desirable in order to However, because engine thrust was not measured, it improve the fidelity of the simul ation for was not possible to identify the important axial force flight-planning purposes. Specialized flight-test parameters. maneuvers (i.e. doublets and POPUs) were flown to support this objective. Analysis of both the short-period (i.e. longitudinal doublet) maneuvers and Analysis of the POPU maneuvers, which span a larger the larger-scale (i. e. POPU) maneuvers allowed range of the flight envelope (airspeed, a, lift coefficient, updating the important parameters of the aerodynamic etc.) allows extraction from flight data of some of the model. Longitudinal stabil ity and control parameters remaining axial force parameters. The POPU were extracted from both the longitudinal doublet and maneuvers were flown with the engine at an idle-thrust the POPU maneuvers using a standard time-domain setting to minimize unknown thrust contributions. output-error parameter estimation code. 17 Hence, the axial-force parameter estimates extracted
3 5 -- Measured 2 4 -- Computed Symmetric a, elevon, 3 deg deg 0 2 -1 20 10
10 Pitch 5 Pitch rate, attitude, deg/sec 0 deg 0
- 10 - 5 1.5 .1 0
Normal Axial acceleration, 1.0 acceleration, .05 9 9
.5 0 2 3 4 0 2 3 4 Time, sec Time, sec 010472
Figure 15 . Comparison of flight-measured and computed time hi story results for a longitudinal doublet in ri gid-wing configurati on.
12 Ameri can Institute of Aeronautics and Astronautics from the POPU maneuvers are more reliable than those curve slope, C N ' was nearly identical to the preflight C< from analysis of the longitudinal doublet maneuvers. prediction. However, the flight-determined estimate of
the longitudinal stability parameter, Cm ' showed lower Figure 16 compares the flight-measured and a stability than the preflight prediction, and the computed time histories for one POPU maneuver analyzed for the preinflated configuration. The large flight-determined estimate of the symmetric elevon excursions in airspeed, a, normal acceleration, and control effectiveness parameter, C ' showed slightly n,6e pitch attitude were expected to be representative of higher effectiveness than the preflight estimate. those observed in the first in-flight inflation of the 1-2000. Table 2. Comparison of analytical and flight-estimated aerodynamic model parameters. Analysis of several longitudinal doublet and POPU Parameter Preflight Flight maneuvers yielded an updated set of longitudinal Prediction Estimate stability and control parameters for u e in the flight (deg-I) (deg-I) planning simulation. Table 2 compares the preflight 0.110 0.105 (analytically-derived) and flight-estimated values of the -0.044 -0.025 primary stability and control parameters--C N ' C m ' a a -0.034 -0.040 C . The flight-determined estimate of normal-force nJ be
Figure 16. Compari son of flight-measured and computed time history results for a POPU maneuver In the preinflated-wing configuration.
13 American In stitute of Aeronautics and Astronauti cs [n-f1ight [nflation the [-2000 (with the wings stowed) coupled with propeller swirl from the carrier aircraft, impart the roll Three flight operations were conducted to rate. Some pitching motion is also apparent in the data. demonstrate the in-flight inflation capability of the 1-2000 and to document the wing and vehicle dynamic Shortly before 27.9 sec, the research pilot response during inflation and transition to lifting flight. commanded wing inflation, and at 27.9 sec, the Figure 17 shows a sequence of six photographic wing-retention straps were released. In figure 17, photo images documenting the air launch of the [-2000 from 3 shows the wing retention straps just after release, the carrier aircraft and the subsequent wing inflation. retracting forward. At 28.05 sec the pressure began to The time listed below each image is only the rise in the wing (fig 18, wing pressure) as it inflated (fig approximate value, estimated by correlating 17, photo 4). As the wing unfolded and inflated, the photographic, video, and onboard measurements, and inertial and aerodynamic effects of the wings generated using a common time scale for all data sources. Figure significant and dynamic rolling moments and heaving 18 shows a time history of some of the pertinent forces, as shown by the roll rate and normal acceleration on board measurements. measurements. Although the photos indicate a Release from the carrier aircraft occurred at 26.9 sec symmetric wing deployment (fig 17, photos 4 and 5), (fig. 18, normal acceleration) at a dynamic pressure of roll rate peaked at -250 deglsec and normal acceleration approxImately. II Ib/ft 2 (fig 18, dynamic pressure), and peaked at 2g. Moments and forces in the remaining the [-2000 was in ballistic flight for about I sec (fig 17, axes were relatively small. The angles of attack and of photos 1 and 2). During ballistic flight the research pilot sideslip induced during the unfolding and inflation were made no control input to the [-2000 beyond the small and well-damped; no indications of divergence or command to initiate the wing inflation sequence. instability were evident. During this time, the dynamic response is primarily a roll to the right (fig 18, roll rate). The low roll inertia of
Figure 17. Photo sequence of air launch and wing inflation.
14 American Institute of Aeronautics and Astronautics Wing retention strap rele7 se Rate 1-2000 release from r Wing retention 1-2000 release from -- Roll carrier aircraft strap release carrier aircraft --- Pitch .- -- - - Yaw 14 100~---+------+------~
I q, Rates, 12 / I -100 I psf ~ deg/sec I -200 r- I I 10 - 300 L...---t----+------..J
-- (J.. --- ~ 200 5r--~---~------~
Wing Flow pressure, 100 Start of I ~ angles, o psig - wing inflation 1/ deg o _5L...----+------+------..J Acceleration .J Differential elevon Normal ___ Symmetric elevon Lateral .- ---- Rudder Axial 5 2r---t----+--;------,
Controls, 1------~---~----:.'..J-- ______Acceleration, deg o g Start of pilot ..J~ o initiated pull-out
I I - 1 L...-__~ ______~ ______-..J -5 26 27 28 29 30 26 27 28 29 30 Time, sec Time, sec 010475
Fi gure 18. Time hi story of air launch and wi ng infl ation.
At approximately 28. 6 sec the wing reached its final Center (Edward s, Californi a). The aircraft was fl own in inflated shape (fi g 17, photo 6) at a wing press ure of a powered configuration with a rigid wing and then with approximately 55 psig. The aerodynami c roll damping an inflatable wing. It was al so air-launched in an of the winged aircraft was now sufficient to damp out all un powered configuration in which th e wing was inflated hi gh-frequency dynamic moti ons. From 28.65 sec in fli ght. Based on the research results to date, the forward , the a and norm al accelerati on time histories following conclu ions have been drawn: show signi ficantly stronger correlation than that during the inflati on process. This is strong ev idence that the I . Recent advances in miniaturized instrumentati on wing is fully inflated and capabl e of generating technology have made it possible to obtain quantitati ve signi ficant aerodynamic lift force. A s the wing pressure fli ght research res ults from aircraft at small scale. continued to ri se toward the fi nal value of 180 psig, the Research data quality was sufficient to allow applicati on research pilot ass umed control of th e aircraft, and fl ew of parameter estimati on fl ight test techniques. the vehicle to an unpowered landing.
2. Mechani cs of materi als analyti cal methods were Conclusions effec ti ve in modeling the multiple-spar wing Ground and fli ght tes t techniques traditionally applied configuration for a ran ge of inflation press ures. to large-scale research aircraft were successfully applied to a small-scale research aircraft config ured with an 3. Integ ration of the inflatabl e wing tes t article into a inflatable wing at th e NASA Dryden Flight Research research aircraft configurati on is poss i bl e at small scale.
15 American Institute of Aeronautics and As tro nautics Powered flight, using only the control surfaces on the http://www.friends-partners.orglmwade/craft/firlide.htm tail of the aircraft, was demonstrated. 7/2101.
4. For the angle-of-attack range spanned in the flight 7U.S. Department of Commerce, Patent and program, the flight data demonstrated the ri gid-wing Trademark Office, Arlington, Virginia, U.S. Patent configuration to be an effective simulator of the 3,944, 169, James R. Bede, Hang Glider, March 16, inflatable-wing configurations. 1976.
8 . . Department of Commerce, Patent and 5. The asymmetric twist distri bution of the inflatable U S Trademark Office, Arlington, Virginia, U.S. Patent wing required significant differential elevon deflection 3,957,232, Wayne A. SebreIl, Inflatable Wing, May 16, to achieve trimmed flight. A small trim tab on one wing 1976. was suffi cient to achieve trimmed flight. 9U.S. Department of Commerce, Patent and 6. The feasibility of ballistic ai rdrop and inflight Trademark Office, Arlington, Virginia, U.S. Patent inflation of the wing, with transition to controll ed lifting 6,082,667, Roy Haggard , Inflated Win gs, July 4, 2000. flight, was demonstrated in three flight operations. Wing inflation and transition to lifting flight was rapid; JOCrimi , Peter, "Divergence of an Inflated Wing," vehicle dynamic response was benign and limited Journal of Aircraft, vol. 37, no. 1, January-February primarily to roll and heave motions. No indications of 2000, pp. 184-186. instability or divergence were evident. J IBrown, Glen, Roy Haggard, and Brook Norton, Acknow ledgment " Inflatable Structures for Deployable Wings," AlAA-2001-2068, A Collection of th e 16th AiAA The authors would like to thank Mr. John Fraysse of Aerodynamic Decelerator Systems Technology the U.S. Navy for making the inflatable wings available Conference and Seminar, Boston, Massachusetts, 21-24 to NASA-Dryden. Without the inflatable wing test May 200 I , pp. 19-26. articles, this resea rch would not have been possible. 12Leonard, Ro bert w. , George W. Brooks and Harvey References G. McComb, Jr., Structural Considerations of Inflatable Reentry Vehicles, NASA TN D-457, September 1960. JU.S. Department of Commerce, Patent and J3Comer, RL. and Samuel Levy, " Deflections of a n Trademark Office, Arlington, Virginia, U.S. Patent, Inflated Circular Cylinder Cantilever Beam." AiAA 1,905,298, Taylor McDaniel, Flying Machine, April 25, Journal, vol. I , no. 7 , July 1963, pp. 1652-1655. 1933.
14Webber, J.P.H ., " Deflections of inflated cylindrical 2S tadvec, Ernest, Th e Inflatoplane, published by cantilever beams subjected to bending and torsion," Essco, Akron, Ohio, February, 1980. Th e Aeronautical Journal, vol. 86, no. 858, October 1982, 3Cocke, Bennie W., Jr. , Wind-Tunn el Investigation of Paper 1020, pp. 306-312. the Aerodynamic and Structural Deflection Characteristics of th e Goodyear Inflatoplan e, J5de Jong, R c., and J. A. Mulder, "Accurate NACA-RM-L58E09, September 10, 1958. Estimation of Aircraft Inertia Characteristics from a Single Suspension Experiment," Journal of Aircraft, 4Whitney, Richard Y. , "Goodyear's Inflatoplane and vol. 24, no. 6, June 1987, pp. 362-370. Project Wagmight," Journal, American Aviation Historical Society, Summer 2000, pp. 104-110. J6Margason, Richard J., a nd John E. Lamar, Vortex-Lattice FORTRAN Program for Estimating 5Jane ,s All th e World's Aircraft, 1955-1956, compiled Subsonic Aerodynamic Characteristics of Complex and edited by Leonard Bridgman, McGraw-Hili, 1956, Planforms, NASA T D-6142, February 1971. p. 90. J7Murray, James E. , and Richard E. Mai ne, pEst 6Wade, Mark, "FIRST Re-Entry Glider," Version 2. 1 Us er's Manual, NASA TM 88280, En cyclopedia of Aeronautics, on the Web at September 1987.
16 American Institute of Aeronautics and Astronautics
------Form Approved REPORT DOCUMENTATION PAGE OMS No. 0704-0188
Public reporting burden for th is collection of Information is estimated to average, hour per response, including the time for reviewing Instructions, searching exiS1lng data sources, gathering and maintaining the data needed. and completing and reViewing the collection of information. Send comments regard ing th is burden estimafe or any other aspect of fh,s collecfion of Information. including suggestions for reducing this burden. to Washington Headquarters Services. Directorate tor Intormation Operations and Reports. 1215 JeHerson Davis Highway. Suite 1204. Arhnglon. VA 22202-4302. and to the Office ot Management and Budget. Paperwork Reduction Project (0704·01 88). Wash ington. DC 20503. 1. AGENCY USE ONLY (Leave blank) 12.REPORT DATE 13. REPORT TYPE AND DATES COVERED January 2002 Technical Memorandum 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS Ground and Right Evaluation of a Small-Scale Inflatable-Winged Aircraft
274-00-00E8-RR-OO-DDF 6. AUTHOR(S) James E. Murray, Joseph W. Pable, Stephen Y. Thornton, Shannon Vogus, Tony Frackowiak, Joe Mello, Brook Norton
7. PERFORMING ORGANIZAnON NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER NASA Dryden Right Research Center PO. Box 273 H-2471 Edwards, California 93523-0273
9. SPONSORINGIMONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORINGIMONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administration NASArrM-2002-21D721 Washington, DC 20546-0001
11. SUPPLEMENTARY NOTES Presented at the 40th AIAA Aerospace Sciences Meeting & Exhibit, 14-17 January 2002, Reno, Nevada.
12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE
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13. ABSTRACT (Maximum 200 words) \ A small-scale, instrumented research aircraft was flown to investigate the flight characteristics of inflatable wings. Ground tests measured the static structural characteristics of the wing at different infl ation pressures, and these results compared favorably with analytical predictions. A research-quality instrumentation system was assembled, largely from commercial off-the-shelf components, and installed in the aircraft. Initial flight operations were conducted with a conventional rigid wing having the same dimensions as the inflatable wing. Subsequent flights were conducted with the inflatable wing. Research maneuvers were executed to identify the trim, aerodynamic performance, and longitudinal stability and control characteristics of the vehicle in its different wing configurations. For the angle-oF-attack range spanned in this flight program, measured flight data demonstrated that the rigid wing was an effective simulator of the lift-generating capability of the inflatable wing. In-flight inflation of the wing was demonstrated in three flight operations, and measured flight data illustrated the dynamic characteristics during wing inflation and transition to controlled lifting flight. Wing inflation was rapid and the vehicle dynamics during inflation and transition were benign. The resulting angles of attack and of sideslip were small , and the dynamic response was limited to roll and heave motions.
14. SUBJECTTERMS 15. NUMBER OF PAGES Research aircraft, Inflatable gliders, Aircraft performance, Inflatable structures, 24 Structural analysis 16. PRICE CODE
17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified U nclassi fied Unclassified Unlimited NSN 7540-01 ·280·5500 Available from the NASA Center for AeroSpace Information, 800 Elkndge Landing Road, Standard Form 298 (Rev. 2-89) Linthicum Heights, MD 21090; (301)621-0390 PrOSCt1bed by ANSI SId . Z39·18 298-102