N93-29715 29 PROJECTARESlI:HIGH-ALTITUDBEATTERY-POWERAEIDRCRAFT CALIFORNSIATATUENIVERSINTOYRTHRIDGE A high-ahimde, battery-powered, propeller-driven aircraft has been designed and is being built by undergraduate students at California State University, Northridge. The aircraft will fly at an altitude of 104,000 ft at Mach 0.2 (190 ft/sec) and will be instrumented to record flight performance data, including low Reynolds number propeller and airfoil information. This project will demonstrate the feasibility of electric-powered flight in a low-density, low-temperature Earth environment that models the atmosphere of Mars. Data collected will be used to design a Mars aircraft to investigate the surface of Mars prior to manned missions. The instrumented payload and the mission profile for the high-altitude Earth flight were determined. Detailed aerodynamic and structural analyses were performed. Control, tracking, and data recording subsystems were developed. Materials were obtained and fabrication begun. The aircraft has a 32-fl wing span, a wing area of 105 sq ft, is 17.5 ft long, has a 12-in payload bay, and weighs 42 lb. It iscomposed primarily of lightweight materials, including Mylar, and composite materials, including graphite/epoxy and aramid core honeycomb sandwich. Low-altitude flight testing to check guidance and control systems and to calibrate data-gathering instruments will take place this summer, followed shortly by the 104,000-ft flight. INTRODUCTION The Universities Space Research Association (USRA), in association with NASA, has sponsored a three-year undergraduate design project in the Mechanical Engineering Department at California State University, Northridge (CSUN). The overall project objective is to design a heavier-than-air craft to fly in the lower martian atmosphere, investigating geological and atmospheric features as a prelude to a manned Mars mission. The first year's design team (ARES I) investigated the martian mission and made recommendations for mission profile, payload, aircraft configuration, and delivery of the craft to Mars. Now, in the second year of the project, a new student team (ARES II) has designed an aircraft to be built and flown on Earth. This prototype, currently being fabricated at CSUN, will i 3 demonstrate the feasibility of the Marscraft by flying on Earth. The aircraft will fly at 104,0OO ft at Mach 0.2, a flight regime closely resembling the martian mission in the following important parameters: atmospheric density and temperature, Reynolds numbers, and lift coefficient. A successful Earth flight will demonstrate the feasibility of the Mars mission and craft. For a meaningful demonstration, the Earth prototype wilt carry an instrument payload to measure and record aircraft performance data. This data will be useful in verifying the design concepts and analytical model developed by the student team and providing a baseline for design of the actual Marscral_, as well as providing valuable low-Reynolds-number propeller and airfoil information. The high-altitude flight is planned for this summer (1991) Phase Description at Edwards Air Force Base. 1-2 Balloon launch and ascent. Release aircraft nosedown at 110,000 ft. MISSION PROFILE 2-3 Pullout to level flight at 104,000 ft. Maximum 3-g load occurs at point 3. 3-4 Preprogrammed level flight and maneuvers for 3minutes. To achieve the stated objectives, the mission profile detailed 4-5 Descend. Wind gusts below 50,000 ft can be dangerous. in Fig. I was developed. A balloon will take the aircraft to 110,000 5-6 Land under manual control. ft and release it. After 3 rain of progranmaed level flight and maneuvers within a 25 by 8 mile test area at Edwards AFB, the aircraft will descend and land under manual control. Fig. 1. Mission profile. 30 Proceedings of the NASA/USRA Advanced Design Program 7th Summer Conference Prior to the high-altitude flight, low-altitude test flights will The wing area was as large as possible to allow for the greatest be made to test guidance and control systems and to calibrate number of solar cells. Ailerons were eliminated for the same data gathering instruments. reason. To simplify fabrication, the wing has a20-ft-span constant chord center section with 6-ft tapered outboard sections. To AERODYNAMIC DESIGN improve stability, the outboard sections have a taper ratio of 7.5, all in the leading edge, to give a sweep effect, and a 13° The basic requirement was to design an aircraft for electrically dihedral, giving the equivalent ofa 5° total wing dihedral. q powered flight at 104,0OO ft (a low-Reynolds-number regime). The tail was designed in conventional configuration, although The selected airfoil, the RG 15-P'I_,is suitable for low Reynolds an inverted V-tail could still be built. The conventional tail is numbers. It has alow design coefficient, small variations in L/D preferred for landing, since the aircraft will not have landing as a function of Reynolds number, and is fabricable due to its gear. Symmetric airfoil sections were selected for both vertical simple shape. The expected Reynolds number at altitude is and horizontal tail sections. Ailerons were not used; instead, 60,000. See Fig. 2 for additional airfoil details. a large rudder and elevator were designed to control pitch, Since the Mars craft will very likely be solar powered, the roll, and yaw. Earth prototype was planned as a solar vehicle too; but due The fuselage was designed in a one-piece "guppy" shape, to cost, availability, and low energy density, solar cells were tapering from a payload bay to a tail boom This configuration waived in favor of nickel-cadmium batteries for this design. simplified structural analysis, eliminated stress concentrations Nevertheless, several key aerodynamic design decisions were associated with a separate boom attachment, and simplified made based on the presence of solar cells, as will be discussed. fabrication. The wetted area was reduced by rounding the square Atrade study was performed based on small payloads, available cross-section, and drag was reduced by giving the payload bay motor horsepower, wing loading vs. velocity, and, finally, wing aslightly sculptured profile. loading vs. aspect ratio. An optimum aspect ratio of 10 was The propulsion system includes an 7-ft propeller, driven by determined. Acarpet plot incorporating these variables resulted a 1.6-hp electric motor through a belt-driven gear reducer, in the simultaneous optimization of wingspan (32 ft), velocity powered by nickel-cadmium batteries (see Fi&3.) With asmaller (190 R/S), and weight (42 lb). Subtracting the known weights payload, or reduced structural weight, the same system can be of the propulsion and avionics systems from this weight resulted powered by K7 solar cells. All the cells can be placed in the in the target structural weight fraction, to be discussed later. wing beneath the clear Mylar skin. Asummary of final aircraft specifications is provided in Fig. 4. .STRUCTURAL DESIGN Large structural elements were divided into finite segments, allowing the analysis of a tapered geometry with nonlinear distribution of loads, moments, and properties. All loads were multiplied by a load factor and safety margin, and the resulting LOW DESIGN COEFFICIENT C(I.)= 0.63 moments used to determine stresses, deflection, and rotation SMALL L/D VARIANCE FOR DIFFERENT REYNOLDS NUMBERS 8.9% THICKNESS SIDLE SHAPE FOR EASE OF FABRICATION I_t_- pT _t'oI) _t.t. t/c =_O_, _ =.tTg.,*t/'c=t.O'.g el _n= tOO,OOO a_ t = l I x./ C_ _f PROPIll.._ O.Ed,._IWO MOTOR 9A'r'r_ ]:1_ • 1 t 1150RPM 8:1RATIO 1.6HI_ la$OO 3.4L8TI-mUg'r BELTDRIVE 1.25I.B I0¢¢U$ 7.0 FTDIAMETER ALUMINUM D,C.DRU_IED 15AH ooo oo_ • e7q o_1 o_ e4_ -io 87%EFFICIENT 2LB 21NX2.75 IN Fig.2. Airfoildata. Fig. 3. Propulsion _hematic. California State University Northrtdge 3 1 PERFORMANCE WING TABLE 2. Final Material Selection Altitude 104,000 ft No sweep Item Material VeI (cruise) Mach 0.02 Wing Area 105 ft^2 VVeell ((cclriumisbe)) 119800 fftt//_se*¢¢ TAasppeerct RRa_aitoio 0.150 Ribs Composite sandwich: 1/4" aramid honeycomb Vel (stall) 172 ft/sec b 32.3 ft core with 7 mil, 1-ply graphite/epoxy (g/e) Load Factor 3 G C (r_,t) 3.5 ft facings Safety Factor 1,2 C (tip) 2.2 ft Torsion box g/e sandwich, allfour sides C Bar 3.34 ft J YBar 7.18 ft Leading and trailing edges 7 mR, 1-ply g/e Skin 0.5 nailMylar, adhered and heat shrunk DESIGN PARAMETERS Fuselage 7 rail, l-ply g/e, with local stiffening Wing Loading 0.4 Ib/ft^2 Weight 42,0 Ib Hp/wt 0,038 For simplification of the wing analysis, it was assumed that Wt/Hp 26.18 the torsion box alone would carry all loads. Normal spanwtse stresses at every box segment were determined using advanced beam theory. The maximum stress in agiven segment was then compared to the local buckling stress, determined from thin- plate buckling theory. Predicted failure stress was kept at least 20% higher than the maximum stress, wing tip deflection was kept less than 1in, and tip rotation was less than 2°. The fuselage was designed similarly to the wing torsion box. VERTICAL TAIL HORIZONTAL TAIL Amostly square cross section was adopted to simplify (1) wing A 1.6 A 6 and tail attachment, (2) analysis and design, and (3) tooling Taper ratio 0.4 Taper ratio 0.4 S 11.89 ft^2 S 18ft^2 and fabrication. The tail section was patterned on the wing. b 4.36 ft b 10.39 ft C (root) 3.89 ft C (root) 2+47ft A number of tests were conducted to provide insight into C (tip) 1.56 ft C (tip) 0,99 ft the behavior and characteristics of certain structural component materials and fabrication methods. A rib test fixture was built Fig. 4, Aircraft specifications. to determine the lowest-weight rib configuration that could sustain the applied load. A thin-film skin test was performed to evaluate skin deflection under load and to observe the behavior of structural assemblies. The geometry, material, or configuration of film and adhesive at low temperatures (-70°F). Lastly, some of each assembly was modified until the computed stresses, informal allowable testing was performed to confirm assump- deflections, and rotations met acceptable limits, and were then tions regarding adhesive and composite strength in tension at refined to reduce weight. low temperatures. The tests were useful in confirming or Table 1 indicates the target weight summary. Note that the modifying the selection of materials. structure comprises 42% of the total aircraft weight, and the A wide range of materials was considered for all structural wing alone weighs O.1 lb/sq ft. Recent advanced aircraft data elements. Primary selection criteria were availability, cost, (e.g., the Gossamer Pengu/n) suggest that these targets are density, and ease of tooling and fabrication. Table 2 indicates achievable. the final materials selection. Figure 5 shows the final con- A maximum load factor of 3 g was determined from an figuration. aerodynamic analysis of the launch. Four-g loads due to gusts may be encountered at 50,000 ft (during descent), determined AVIONICS from velocity-load diagrams. Nevertheless, it was decided to use a design load factor of 3 Pvto help reduce weight, and fly on Systems for flight control and tracking, testing, and data re- a low-gust day. Similarly, a safety factor of 1.2 was selected, cording were designed, tested, and assembled. Flight control even though 1.5 is typical in aircraft applications. will be provided via modified radio-controlled model aircraft hardware. The transmitter station will include an amplifier and directional antenna to boost the control signal to the required TABLE1. Weight Summary range. The aircraft will carry a miniature radio transponder to aid in radar tracking. The control system incorporates a fail- Item Weight (lb) safe feature whereby the aircraft assumes a preprograrmned Structure W'mg 10.5 descent profile in the event of aloss of control signal. Fuselage 12.0 The aircraft will carry sensors to measure and record airspeed, Tail 3.0 altitude, ambient temperature, motor power consumption, and Propulsion propeller thrust and speed. This information will be stored for Propeller 4.4 later retrieval by the flight data recorder (FDR). The FDR uses Motor 1.6 Gear reducer 2.0 the latest in LSI and surface-mourn chip technology to pack Batteries 2.5 what is,in essence, acomplete microcomputer onto afew square r_oad 6.0 inches. The FDR will also provide sensor sigmd conditioning Total 42.0 and supply regulated reference voltages. 32 Proceedings of the NASA/USRA Advanced Design Program 7th Summer Conference q Fig.5. Configuration drawing. Additionally, the aircraft ,_dl/ carry a wing pressure port TABLE3. Budget scanner system to record air pressure differentials at 32 different Item Cost, $ chord-wise wing locations. This data will validate the comput_nalfluid dynamics model. Avionics 8,600 Because of low ambient temperatures at altitude (-70°F), Propulsion 3,800 Structure 19,000 the instrumentation package will be irxsulated. The FDR will Mission Support 5,300 be able to monitor and regulate the payload bay temperature Travel 5,800 with the aid of a small film resistance heater. Power for the Total 42,500 onboard avionics will be provided by a dedicated carbonmo- noi_uoride Iithium battery. ThOrFIDI_will be furnlshed independ- ently with a battery backup. Figure 6 illustrates the control, .i data collect|on, and flight data recorder schematics. Most of the structural components were made from fiber/ epoxy sheets or composite sandwich (an aramid honeycomb BUDGET core with graphite/el_xy facings). Molds, made of wood or sheet metal, were built, surfaced with putty and enamel paint, In a project of this scope, the required instrumentation and then prepared with release agents to accept the wetted cloth. materiel support is substantial. "_anks to the generous donations Curing was done in vacuum. A brief discussion of fabrication of interested industrial associates, the fabrication phase has methods used for the major components follows. proceeded on schedule with a continuing promise of success. The wing isbeing built in three sections (a 20-ft center section Table 3 is abudget of this year's work and two dihedral sections), due to fabrication space limitations and t_rtation di_culttes. The wing torsion box for each FABRICATION section was laid up full length in two L-shaped pieces and assembled with epoxy. Mylar tape was used to attach the ribs The aircraft is being constructed entirely by the student team. to the box, and to face the edges of the ribs for skin attachment. in.dual elements and major subassemblies are being built The leading and trailing edge caps were made in 4-ft sections and tested, and redesigned ifnecessary, before final construction, and adhered to the ribs. The Mylar skin, in 50-in wide strips, California State University Northridge 33 Controls Schematic tail, motor, and payload mounting. The two halves were attached with adhesive tape and epoxy at the seam and around the bulkheads. Tail assemblies were built similarly to the wing. Aluminum brackets were made to secure the wing and tails to the fuselage and permit disassembly for transport. CONCLUSION Through the successful flight of ahigh-altitude Earth prototype aircraft, CSUN Mechanical Engineering students hope to verify the feasibility of a heavier-than-air craft for the martian atmo- sphere. Though the prototype isbattery-powered, it is reasonable to expect the Marscraft to be solar-powered. Some of the major design challenges included (1) using low Reynolds number airfoils, (2) electric motor selection and power system development, (3) low Reynolds number propeller design, (4) structural analysis of composite aircraft components, Paylcad?5cnsors and (5) high-altitude, low-density atmospheric flight perform- s,,,¢ Air aIlce. The data collected from various onboard sensors will be used to build a database that can be used to improve the aircraft design for subsequent flights. Next year's Project Ares III will benefit from our experience in aircraft structural design and ILtmAw_ i_ fabrication, in high-efficiency aerodynamic design (verified by CFD), and in aircraft guidance and control with onboard data acquisition. With this knowledge, future design objectives can include detailed aircraft subsystems such as flaps, slots, landing gear, video cameras, autonomous control systems, and refined aerodynamic and structural designs. The Ares IIdesign team has responded to adifficult challenge: the design and fabrication of an all-composite, high-attitude, electrically powered aircraft. In the summer of 1991, we will fly this aircraft at an altitude of 104,000 ft. This accomplishment will generate world-wide interest and set CSUN and its students _aU.l_na.ilma_ apart for having flown at an altitude never yet achieved by _E"T T_ ulednfV propeller-driven aircraft. tt t! ACKNO_LEDGMENTS r The project Ares IIconsists of EdAvetesian, Tony Banh, Doug Boobar, Chris Charan, Jerry Gaudreau, Scott Gordon, Elise 11 11 Gravance, David Hooper, Elizabeth Kralian, Bogdan Macri, Denise Mah, Doug Meyer, Margie Monsen, Craig Morton, Fred Mugford, Mel Navarro, Ermin Nazareno, George Renteria, Richard Shafiroff, Ralph Valencia, and Tammy Yeast. ._m_Caml_ IIA_lOll• 17DigitaClMm_b Sql_JC_il_q 2M_Dtu_r_t De-,S_nlff I0lh;J_'loluum PowtrC_itioaiat IL_232C_l_ubk ruttutlt L*t_ laura'Sla,ck. Fig.6. Avionics schematics. was adhered chordwise to the wing skeleton and heat shrunk. Connectors were made to join the 6-ft dihedral sections to the center wing section. The fuselage has a symmetrical cross-section, so the top and bottom halves were made full-length fi'om the same mold. Bulkheads were installed for shape control and for the wing,