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JOINT USAF/NASA HYPERSONIC RESEARCH AIRCRAFT STUDY

Frank S. Kirkham and Robert A. Jones NASA Langley Research Center

Hampton, Virginia

and

Melvin L. Buck and William P. Zima Air Force Flight Dynamics Laboratory

Wright-Patterson Air Force Base, Ohio

Presented at the AIAA Aircraft Systems and Technology Meeting

Los Angeles, California August 4-8, 1975

JOINT USAF/NASA HYPERSONIC RESEARCH AIRCRAFT STUDY

F. S. Kirkham* and R. A. Jones** NASA Langley Research Center

Hampton, Virginia and

M. L. Buck+ and W. P. Zima++ Air Force Flight Dynamics Laboratory. Wright-Patterson Air Force Base, Ohio

Abstract

A joint USAF/NASA study has developed a con­ceptual design for a new high-speed research airplane (X-24C) and identified candidate flight research experiments in the Mach 3 to 6 speed range. Four major categories of high priority research experiments are described as well as the X-24C design concept. The vehicle, a rocket-boosted, delta planform aircraft, is air launched from a B-52 and is capable of forty seconds of rocket cruise at Mach 6 with a research scramjet. Re­search provisions include a cedicated 10-foot long research experiments section, removable fins and stra~~s, and provisions for testing integrated airbredthing propulsion systems.

Introduction

Advanced operational airbreathing aircraft have been limited by aerodynamic heating and turbo­jet engine perfonnance to speeds of about 11ach 3. It is now certain, however, that a wide variety of potentially useful and effective aircraft can be built to operate at much higher speeds.

The systems which appear most promising include hydrogen-fueled airbreathing transports of potential interest for very long ranges at speeds of Mach 4 to 8. The airbreathing launch vehicle (Mach 7 to 10) is an aircraft-type first stage candidate for future space shuttle systems which offers not only reusability but many operational advantages. Other airbreathing systems for possible future military applications include advanced reconnais­sance and fighter/interceptor-type aircraft, and high-speed strategic systems. A postulated weapon, the ram-airbreathing laser, would also operate at these speeds (Mach 3 to 6).

During the past 15 years considerable progress has been made in ground-based R&D aimed at solving the critical problems of future high-speed aircraft. Many of these developments have reached the stage where they must be demonstrated in actual flight at large scale prior to applications. Aeronautical history for all vehicle types reveals without ex­ception that technology development has required involvement with a manned flight vehicle to realize the ultimate goal of technological maturity and' useful applications. The research airplane provides focus and stimulus for ground-based research and development and demands a level of commitment which will guarantee flight worthy hardware. While it is

evident that much of the value of a flight test pro­gram is derived from the focused laboratory develop­ment and testing which it generates, the actual flight demonstration in the real environment of large critical components, such as, advanced struc­ture or a new propulsion system enables decision makers to accept these technologies as proven options for future operational systems.

Specifically, the principal goal is to develop a new high-speed research airplane capable of pro­viding:

Evaluation of aerodynamics, stability, and handling qualities of blended wing body aircraft.

Demonstration of advanced airbreathing· propulsion systems and their integration with the airframe under conditions not reproducible in ground facilities. The propulsion systems would include:

Ramjet Scramj et Composite engines

Demonstration of advanced structural system assemblies such as:

Hot structures Insulated structures Actively cooled structures LH2 tankage

Tests of miscellaneous components 'and sub-systems such as:

Sensors, optics Radomes Laser systems Weapons

Previous Research Aircraft

The X-15 rocket research airplane perfonned valuable pioneering flight explorations at speeds up to Mach 6.7 which provided an essential back­ground for many of the foreseen high-speed aircraft applications. Among the X-15's contributions were verification of the nature of the hypersonic heat­ing environment. validation of wind tunnel tech­niques, establishment of piloting and energy management techniques for lifting reentry

* Head. Hypersonic Aerodynamics Branch, HSAD, Member AIAA **Head. Hypersonic Propulsion Branch, HSAD, Member AIAA + Head, High-Speed Aero Perfonnance Branch, Member AIAA ++Technical Manager, Vehicle Applications Group, Member AIAA

from space, and the performance of other s pace­related experiments. However, the X-15 was not equipped to demonstrate any of the new technologies in airbreathing propulsion, hypersonic structures, or blended wing-body aerodynamics which require flight testing before they can be applied to future operational aircraft.

Since the termination of the X-1S program in 1968, several proposals for new high-speed research airplanes have been made. Illustrations from some of the more recent studies are given in figure 1. These manned research vehicles can be grouped as fo11 ows: (1) Advanced vehi c 1 es i ncorpora ti ng the currently-favored new structural, propulsive, and configurational features; and (2) Near state-of-the­art type vehicles capable of accepting a wide variety of large research experiment payloads, and capable of relatively low-cost modifications for performance growth.

Those research aircraft considered in the first group were often aimed specifically at a parti­cular class of application. Recent examples would be the NASA Hypersonic Research Facilities Study (HYFAC) proposal, the NASA Langley Hypersonic Research Aircraft (HRA), and the Air Force's Mach 3 to S Flight Test Vehicle. Aircraft in this class were large vehicles sometimes requiring new launch systems; they were very expensive; and they were inflexible in that they represented only one type of future aircraft with a particular set of ad­vanced technology requirements. Since ideas as to the best technological approaches may change, vehicles in this group could possibly become pre­maturely obsolete.

The second group of vehicles is made up of smaller aircraft with reduced vehicle and program costs. In this concept, the basic aircraft structural and propulsion systems are conventional or near-state-of-the-art. To minimize initial cost, advanced technology demonstrations are con­sidered only as research payloads. Examples of vehicles in this group for which studies were recently completed are the NASA-Langley High-Speed Research Aircraft (HSRA) and the Air Force Flight Dynamics Laboratory Incremental Growth Vehicle (IGV). Both of these concepts used the existing B-S2 launch system (used by the X-1S, X-24A, and X-24B) and offered considerable flexibility to accommodate a variety of experimental research payloads. While their flexibility made these concepts more cost effective than the vehicles in the first group, these aircraft were still rela­tively expensive.

In addition to these major high-speed research aircraft systems a number of special purpose air­craft with more limited objectives have been built. Examples are the USAF X-24A and NASA HL-10 and M2F2 programs designed specifically to explore the piloting problems of lifting-body reentry vehicles at low speeds and landing. The successful X-24. program was extended by USAF to include a more slender shape, the X-24B, at speeds up to about Mach 2. Recently the Air Force proposed an additional low cost extension to about Mach 5, denoted as the X-24C.

Obviously, such a vehicle could accomplish some of the objectives listed above for the high speed research airplane systems, and the question arose as whether some other low-cost derivative of

2

the X-24 configuration could be designed to accom­plish all of the major objectives of a new high­speed research airplane.

Since considerable interest was evident on the part of both USAF and NASA an ad hoc study group was formed in May 1974. The study centered on the use of the X-24C concept to develop a flight vehicle which would fulfill USAF and NASA research objectives.

Vehicle Requirements

The approach taken was to develop a flight research vehicle which has the inherent capability to be a test bed for a significant number of experiments -- not only those we can visualize now but the unknown experiments of the future. Since the design requirements of the vehicle are strongly influenced by the experiment objectives the first task was to define the critical technology areas and to determine flight requirements.

Potential experiments were first screened to identify those of outstanding benefit to future operational systems and then separated into two categories: those which have a major impact on the research vehicle design (or performance), and those which have only a minor impact. For those experiments having a major impact on the vehicle design, the minimum requirements for flight test in terms of size, weight, configuration, speed, altitude, test time, instrumentation, and number of flights were selected on the basis of best collective judgement of the study group. Incre­ments above the minimum requirements were also considered in terms of increased research value. Relative costs and research value of vehicle and candidate experiments were considered in develop­ing a cost effective set of research vehicle requi rements.

Candidate Flight Experiments

Three major classes of experiments -- Configu­ration Technology, PropUlsion, and Structures and Thermal Protection Systems -- were selected as representing the most demanding requirements for flight research vehicle design and were used to define the configuration, performance, and research payload size, volume, and weight. Examples of these major experiments which drive vehicle design requirements are:

Configuration Technology: A representative list of some of the important technical disciplines in which tests of the blended wing body configura­tion will contribute to future military and civil flight vehicles is given in figure 2. The X-24C configuration has a general similarity to possible future systems as shown. The aerodynamics, hand­ling qualities, stability and control, etc. dis­ciplines shown require flight demonstration in a manned aircraft before application to operational systems.

Propulsion: As pointed out by the Supersonic Technology subpanel of the Joint DOD/NJ',SA Aero­nautical Research and Development Study, there are technology developments on the horizon that hold potential promise for significantly improving future military vehicle capability. The choice of engine cycle for these future systems may vary greatly from one vehicle to another as mission and

operating requirements dictate. Turbojet/ramjet systems will be a strong contender for some types of interceptors. For other mission requirements novel engine concepts such as the supercharged ejector ramjet are of interest. For missiles some combination of rocket and rarrjet seems most appropriate.

Studies indicate that the integrated, hydrogen­fueled scramjet is the most promising air-breathing propulsion system for speeds above Mach 6. Not only does the scramjet provide higher specific im­pulse than the ramjet at these speeds, as shown in figure 3, but the cryogenic hydrogen fuel can be used as a heat sink for cooling the engine struc­ture and a part or all of the airframe. The fixed-geometry, modular scramjet currently under development at NASA-Langley is a principal can­didate flight demonstration and was thus used in defining propulsion flight test requirements. Subsonic combustion ramjets were also examined and flight test requirements found to be within the envelope of requirements established for the scram­jet.

I twas determi ned tha t f~ach 6 is the lower limit for adequate demonstration of the airframe­integrated scramjet (figure 4). Since no clean air facilities presently exist for propulsion testing at Mach numbers above 5, flight tests are required for verification of combustion aerothermodynamics. In addition, the limited size of existing facili­ties makes flight testing a necessity to confirm the interactions of propulsion system with the airframe and define the contribution of the after­body/nozzle combination to thrust, lift, and pitching moment. At least three modules are required to determine structural and aerodynamic interactions between modules. Further, each scramjet module must be at least l6-inches high to allow construction of a representative, regener­ative1y cooled structure. The vehicle lower sur­faces must also be configured to provide inlet precompression, the afterbody designed to serve as an exhaust nozzle, and payload volume provided for the hydrogen fuel and subsystems. A minimum sustained flight time of 40 seconds at Mach 6 at a dynamic pressure of 1000 1b/ft2 is required to allow vehicle stabilization, starting the engine, retrim of the aircraft, and data acquisition. This would require sustained rocket thrust to supplement the thrust of the three scramjetmodu1es.

A further valuable flight test would be the demonstration of cruise on scramjet power alone. The requirement of Mach 5 plus 40 seconds of rocket cruise with three 16-inch scramjet modules will also provide the capability to take six larger scramjet modules capable of cruise thrust to Mach 6 at rocket burnout thus allowing demon­stration of cruise on scramjet power alone.

Structures and Thermal Protection Srstems: As with other technology demonstrations, aarge part of the value of structural flight projects is derived from laboratory development and ground tests which precede the flights. However, it 1s the existance of the flight vehicle which provides meaningfully detailed design criteria, requirements and justifications without which major structural ground-based developments are seldom undertaken. Actual flight demonstrations in a manned aircraft of large critical portions of advanced structure

3

will substantially reduce risk inherent in commiting to major programs requiring advanced state-of-the­art structural solutions. To demonstrate confidence, the experimental structure must be sufficiently large to realistically duplicate the operational vehicle application with all of the principal environmental factors present. The fact that flight test struc­tures designed with normal margins of safety perform successfully provides confidence in the concept for future application.

Provisions must be made to test both advanced fuselage and empennage structures. Examples of such advanced structures are: hot structure, insulated structure, actively cooled structure, integral liquid hydrogen tanks and large areas of thermal protection. A replaceable fuselage section approxi­mately two diameters long (about 10 feet for a 50 foot vehicle) provides capability for adequate test of major fuselage structural components (figure 5). Transition structures to take the loads from the experimental payload to the baseline aircraft would be included as a part of the research experiment as shown in the integral liquid hydrogen tank payload. The replaceable fuselage section of the baseline vehicle includes a stand-off TPS designed to allow test of a variety of materials and TPS concepts. The heat flux distribution around the fuselage payload section (shown in figure 5 for 11ach 6) pro­vides a realistic environment for test gaps, ex­pansion joints, seals, and thermal distortion problems assiciated with hot structures or metallic TPS systems.

An additional requirement for flight test of structures and thermal protection systems is replaceable fins and strakes with provisions to accommodate the differences in thermal expansion between the fin and fuselage structure. Such a provision allows test of lifting surface structure, leading edges and surfaces subject to interference heating. The active cooling ·experiment shown in figure 6 illustrates one possibility that employs a replaceable vertical tail.

Other Ex~eriments: The versatile research approach outllned above affords numerous additional experimental capabilities (figure 7). For example, the integral rocket-ramjet which has potential application for long range, high Mach number cruise missiles as well as inlets for subsonic combustion ramjets have been identified as potential candidates for flight demonstration. Two types of subsonic combustion ramjets are of principal interest: one which provides low drag and good performance during acceleration at some sacrifice in cruise performance, or a high efficiency cruise system which may incur off-design penalties. Flight test offers the oppor­tunity to verify wind tunnel to flight correlation parameters and scaling procedures and the ability to determine inlet/engine interactions which cannot be adequately defined in ground test facilities.

The large, interchangeable payload bay and replaceable strakes and fins afford an opportunity to demonstrate in flight other advanced technology disciplines of importance to future military appli­cations. Among these are: radomes, antennas, radar absorbing materials, infra-red suppression tech­niques, optical systems, advanced sensors, and weapons separation and delivery. Other experiments such as the determination of boundary layer tran­sition and noise, interference heating, control

effectiveness, etc. have application to both civil and military aircraft. Further, the development of a versatile research aircraft will stimulate new ideas not yet identified as candidates for flight demonstration. One example might be a ram-airbreathing laser weapon.

Vehicle Design

The original research vehicle concept adopted as a starting point for this study (figure 8) was primarily intended to explore the aerodynamic and heating characteristics of a blended wing-body delta-planform aircraft in the Mach 2 to 5 speed regime, and was conceived as a growth version of the X-24B. Maximum emphasis was placed on achieving minimum initial cost. Accordingly, this original concept used many of the components and parts from the X-24B, was air launched from the B-52, used the XLR-99 rocket engine from the X-15, and pro­posed an insulator/ablator thermal protection system directly bonded to an aluminum primary structure. A number of fluid mechanics and other research experiments were suggested for the basic vehicle with its small, dedicated payload bay(s). However, no provisions were made for large scale systems demonstrations of advanced propulsion, structural, and LH2 tankage systems under devel­opment in NASA/LRC programs.

The joint NASA/USAF flight experiment de­finition studies described previously resulted in requirements for higher speeds and the ability to accommodate payloads of greater weight and larger volume than originally envisioned. A series of trade studies were conducted to develop a vehicle concept which provided maximum research versatility at minimum cost. Alternate rocket propulsion systems were explored as one means of achieving increased performance. Further, the necessity for providing a large volume payload section without compromising performance led to an investigation of propellant tank concepts with high volumetric efficiency. Extensive trade studies of potential structural/thermal protection system solutions were conducted to define a concept capable of routinely carrying flight experiments into the Mach 6+ speed range while keeping vehicle turn­around time and operational costs within reason.

Configuration Modifications: As shown in figure 9, the first vehicle modification was primarily aimed at increasing maximum speed capa­bility. Vehicle size was geometrically increased and the propellant tanks were lengthened. However, the vehicle did not meet the 40-second rocket cruise criteria at Mach 6 with a research scramjet.

The second vehicle modification (figure 10) increased performance through increased propellant volume with no further change in the length or planform of the vehicle. The cockpit was moved forward and the body depth increased to enlarge propellant volume and improve the bottom surface geometry for airbreathing propulsion tests. Al­though the vehicle performance was thus increased to 17 seconds of rocket cruise at Mach 6 with the scramjet payload, no provisions were made for a dedicated body section for structure and hydrogen tank tests.

Propellant Tank and Rocket Trade: At this point in the conflguration evolution, propellant tank trades combined with configuration modifi-

4

cations were studied in order to provide a dedi­cated fuselage section for major structural tests without sacrificing propellant volume and perfor­mance. As shown in figure 11, the lower surface ramp was moved further aft to improve conditions for the scramjet experiment and the cockpit was moved further forward to provide additional pro­pellant volume. To maintain the propellant volume of the original propellant tanks an inter­secting lobe tank with a honeycomb sandwich common bulkhead was used. These changes resulted in a ten-foot long payload bay free of basic propulsion tanks. This payload section is removable via built-in field splices such that future experimen­tal sections can be inserted for flight tests.

One obvious way to achieve 40 seconds of quasi-steady-state rocket cruise at Mach 6 with a small volume-constrained research airplane would be to develop a new high-pressure rocket engine. However, to minimize cost, only existing rocket engines were considered. A preliminary screening revealed that the five rocket systems illustrated in figure 17 had characteristics consistent with performance requirements. These systems were examined in more depth using cost, performance, and environmental· impact as the selection criteria. Although high density storable propellants pro­vide superior performance, the LR-Sl (Agena) engine was eliminated because of the highly toxic pro­pellants. The RL-10 (Centaur), a high specific impulse, hydrogen-fueled rocket was rejected because the low density propellants increased the vehicle size and thus cost (for constant perfor­mance). Thus, the LOX/NH3 rocket used in the X-15 program (XLR-99) and two LOX/RP rockets used in the manned space program were selected as final candidates. The LR-9l engine is available from government storage and has roughly the same

. performance characteristics as the LR-10S. For the purposes of this study, the XLR-99 was selected as the baseline engine and the LR-105 engine was considered the primary back-up engine.

The recommended propulsion system candidates are shown in figure 13. The X-15 engine (XLR-99) with an extended nozzle for improved performance has been selected as the baseline rocket propulsion system based on the experience accrued wMth this engine by AFFTC/FRC during the X-1S program, proven operation in a research airplane program, throttle capability to 40-percent of maximum thrust and the availability of six engines and parts capable of servicing a 100-flight program.

LR-10S engine is man-rated and has established a good reliability record as the Atlas sustainer engine. However, engine modifications would be required to provide horizontal starting, throttle capability for more versatile flight operations, and qualifications for starting at altitude. If any problems result from the use of the XLR-99 engine, the LR-10S rocket is the back-up system.

Using these results, a third configuration modification, shown in figure 14, was devised. Lobed nonintegral propellant tanks used in combi­nation with the XLR-99 rocket with an extended nozzle resulted in a configuration concept meeting the established performance criteria.

Structure/TPS Trade: The most difficult pro­blem in designing a low cost, Mach 6+, research

vehicle is developing a reliable. low-cost thermal­structural approach. Several alternative systems shown in figure lS were investigated. The relative merits of these systems are:

(1) An ablator directly bonded to an aluminum structure can potentially act as an insulator allowing repeated flights without refurbishment -- provided that the pyrolized surface formed in flight is retained. Initial cost and weight are lower than other systems which could have a sig­nificant impact on minimizing vehicle and sub­systems costs. Another bond-on candidate, reusable surface insulation (RSI) being developed for the Space Shuttle can be considered as a back-up with somewhat higher cost than the insulator/ablator.

(2) Hot structures. used for the XB-70 and SR-7l offer a very serviceable surface. but at greater weight and substantially higher initial cost than the direct bond insulator/ablator system. The hot (1200°F) inconel heat sink structure used for the X-1S was also found to be more expensive than other reusable metallic TPS. For this reason these types of structures have been eliminated.

(3) Carrier panels supporting insulator/ ablator or RSI provide research versatility by enabling tests of future TPS panels at any lo­cation on the airplane. However, carrier panel weight and cost are disadvantages. Insulator/ ablator or RSI on carrier panels remain as candi­dates for the payload bay for research versatility.

(4) Although metallic radiative systems have the potential for long life at low weight, the combination of a lightweight heat shield and packaged insulation is more complex and expensive than other competing systems for the research vehicle and was thus dropped.

(5) Using stand-off shields of sufficient thickness to serve as a heat sink to limit surface temperature to below 600°F avoids some of the problems of radiative shield systems. Although aluminum heat sink shields were too heavy for this application and therefore were eliminated, heat sink shields constructed from a beryllium-aluminum (Be-38Al) composite remain as a candidate TPS for the payload section.

(6) A heat sink structure combines load­bearing and thermal protection functions thereby reducing the combined weight of the structure and thermal protection from that of heat sink shields on standoffs. The Be-38Al heat sink structure is no more than l5-percent heavier than the com­bined weight of the insulator/ablator and aluminum structure. This weight increase may require either added tank volume or a more efficient rocket to achieve the performance goal. However, a heat sink structure provides a very durable vehicle requiring no refurbishment. Cost estimates have shown the heat sink structure to have a somewhat greater' initial cost than the insulator/ablator system. but the initial cost may be quickly balanced by ablator refurbishment costs. For these reasons the Be-38Al heat sink structure is a contender for the basic research vehicle structure/TPS.

The final candidate thermal protection systems are shown in figure 16. One system is an external insulator/ablator bonded to the skin of the aluminum load-carrying structure. with a 2S0°F

S

aluminum structure temperature limit. An alternate system is a heat sink load-carrying structure of Be-38Al with a 600°F structural limit.

The advantages of the bond-on approach propos­ed in preliminary evaluations are its initial low

. cost. low weight, potential for reuse at Mach 6 and its compatibility with a conventional aluminum airframe.

The advantages of the heat sink system are primarily its durability, with virtually infinite reuse. the capability of providing almost unlimited internal access through a bolt-on skin system, potentially competitive initial cost, and the ability to change selected heat shields to accommodate increased vehicle performance.

These two thermal protection systems offer completely different solutions to the problem and each will be carefully evaluated before a final selection is made.

Proposed USAF-NASA High-Speed Research Aircraft

A general arrangement of the proposed USAF­NASA research vehicle is shown in figure 17. The configuration. a generic variation of the X-24B aircraft. is a 48.S foot long. blended wing­body vehicle of delta planform. The forebody. is tailored to provide acceptable flow atthe inlet face for research propulsion systems mounted on the fuselage lower surface. The afterbody serves as an exhaust nozzle for integrated research scram­jet packages. The vertical tails are sized for directional stability throughout the flight enve­lope. The center tail fits through the cutout in the trailing edge of the B-S2 wing when the vehicle is mounted on the launch aircraft. Speed brakes are being considered on the center fin as well as in other locations.

As shown in figure 18, the cockpit and most non-research equipment are located in the forward fuselage. The large, interchangeable flight re­search experiment section located aft of the cock­pit area is reserved for flight experiments. The use of a fly-by-wire control system mini'mizes the routing of mechanical systems through the payload section and thus facilitates interchangeability for a variety of payloads. The aft fuselage con­tains the rocket propellant tanks. main landing gear. and the rocket primary propulsion system. The XLR-99 rocket engine with an extended nozzle has been selected as the baseline engine. The LR-10S is the back-up rocket.

A major goal of this program will be to keep the research vehicle acquisition cost minimal. One major factor in minimizing costs is the use of government-furnished equipment (GFE) available from current stocks. Among the GFE subsystems are the rocket engine. the ejection seat, landing gear. communication system, and approximately SO other smaller items. Demonstrated advanced tech­nologies from past (eg. X-1S) and present programs must be utilized to the fullest. Included in this category are such items as auxiliary power units and/or batteries, inertial and/or air flight data systems, research data recording systems and flight control systems components such as adaptive control and fly-by-wire technology from such pro­grams as YF-16. Another element in maintaining a

low cost program is the use of the existing High Range at Edwards Air Force Base. The flight trajectories and speed brakes of the proposed vehicle were specifically tailored to meet this requirement.

As illustrated in figure 19, the proposed research vehicle meets the flight research require­ments. ~linimum design performance is 40-seconds of rocket cruise at ~lach 6 with a payload drag and weight equivalent to three l6-inch scramjet modules. The vehicle is designed for a maximum dynamic pressure of 1250 psf but will operate from a dynamic pressure of 1000 psf to about 50 psf. In addition to acquiring aerodynamic and handling quality data, the flight test results will contri­bute to an understanding and definition of opera­tional problems such as maneuvering and tracking, radar signature, and infra-red signature. De­finition of flight sensors and displays for typical missions coupled with the man-aircraft interface will provide information not available through simulation. The development of these data and the verification of both analytical and ground­based experimental results is of the utmost im­portance to establish confidence for the definition of future military systems. The basic aircraft will be fully instrumented to measure these characteristics and provisions for other experi­ments such as boundary layer survey rakes and interference heating calorimeters will also be included.

The bottom surface of the vehicle is config­ured to accept research ramjet or scramjet systems integrated with the airframe in a manner consistent with future applications. Fuel and instrumentation for the propulsion tests will be carried in the payload section. The ten-foot long payload section provides the flexibility to conduct numerous flight experiments such as weapons de­livery or avionics tests. Large-scale structural demonstrations can be conducted by replacing the entire payload section with a test structure such as an integral hydrogen tank, an actively cooled airframe section, a hot structure or an insulated structure. Thermal protection systems can be demonstrated by replacing part or all of the stand-off TPS on the payload section. Further­more, advanced structures may be demonstrated by replacing the strakes or fins.

Growth Potential

The ultimate performance capabilities of the proposed research airplane will largely depend on the final selection of the thermal protection system and rocket. Extensive testing and develop­ment programs and preliminary design activities are planned to determine the utility and cost effectiveness of the two thermal protection system candidates prior to vehicle acquisition. The insulator/ablator and XLR-99 approach may prove to be limited to the desired 40 seconds rocket. cruise at Mach 6; however, present preliminary test results indicate that higher speed performance may be possible without the need for significant thermal protection system refurbishment. If, however, the heat sink structure with the LR-105 rocket approach is proven to be most cost effective, it affords the potential for substantially higher performance than' has heretofore been discussed. In the case of both thermal protection system/ structure concepts, two buJlt-in factors afford

6

the potential for increased speed performance: (1) the propellants w~ich are used to "cruise" with 40 seconds can be used instead to achieve greater peak velocities and (2) the lO-foot payload-bay section in the vehicle can be used to store more rocket propellant.

HRA - 197) ISUBSCALE PROTOTYPE I

M·lD

HSRA-1974 ISYSTEMS DEMONSTRATOR I

~ PRESENT STUDY 1975

NASA STUDIES _________

d USAF STUDIES----

/}...­~

M,)- 5 DEMONSTRATOR - 1972

ISUBSCALE PROTOTYPE)

IGV - 197J IGROWTH VEH I CLE) M'4,5-6-9

~ ORIGINAL X-24C M; 5

Figure 1.- Recent research aircraft concepts.

[>-C;:0RBITAL VEHICLE

~ • AERODYNAMI CiHEATI NG/DYNAMI C

CHARACTERISTICS

• STABILITY AND FLIGHT CONTROL SYSTEMS

• MAN/AIRCRAFT INTERFACE HIGH-SPEED CRUISE VEHICLE

~ Ats=ERCEPTOR AIRCRAFT

• MANEUVERING AND TRACKING

• IR/RADAR SIGNATURE

• FLIGHT SENSORS AND DISPLAYS

• CONTROL SURFACE SEALS

MACH) TO 5 DEMONSTRATOR

Figure 2.- Vehicle technology requirement - a blended wing-body configuration.

RELATIVE SPECIFIC IMPULSE

Figure 3.- Relative performance of hydrogen­fueled air-breathing propulsion systems.

REQU I REMENTS PAYOFF

• MACH 6 - 40 sec CRUISE - lOOOQ • DEMONSTRATED PERFORMANCE OF AIR-

• HZ TANK AND SYSTEMS BREATHING HYPERSONIC PROPULSION

• ACCOMODATE 3 MODULE INTEGRATED SCRAMJET

• AERO CONTROL TO HANDLE ENGINE­AIRFRAME INTERACTIONS

• 300 - 1000 Ib/ft2 DYNAMI C PRESSURE

HYDROGEN FUEL REGENERATIVE COOLING THRUST VEHiClE INTEGRATION

• AERODYNAMIC INTERACTIONS

• OPERATIONAL ASPECTS

Figure 4.- Flight test requirements for hyper­sonic propulsion systems.

H[ATING DI SIRI BUlION.

BTU/tt2 sec REQUIR1"IN1S

• REP1AC1ABL[ FUSElAGI I[ClIO'I. 10 Fm LONG • HYP[RSG:-.I C HEAT\~,G A/.D lOADS • QUASI SHADY 1[\1 m',E • \[HIClE l\rERFAC[S

ID

2.5

Z 77

PAYOff r--5"

EXM.1PLE, ~. DEMO~STRA1ED SIRUCTLRAL lICHNOLOGIES

INTEGRAL LH2 lANK : ~~~u~~~~gT~TR:~CTURES • I NTEGRAL 1HZ TANK S

• ~[TA,lIC IPS • SEALS GAPS. EXPANSION JOINTS

• FLIGHT OP[RATIONAL EXPERIENCE • GROUND INSP[CTION AND MAINTENANC[

Figure 5.- Flight test requirements for fuselage structures.

REQUI REMENTS EXAMPLE,

• REPLACEABLE FINS AND STRAKES ACTI VE COOLI ~G

• VOLUME FOR SUBSYSTEMS • VEHICLE INTERFACES • HYPERSONIC HEATING AND LOADS .QUASI STEADY TEST TIME

PAY-Off

• LEADING EDGE, SEAL, ACREAGE DESIGN • LOW WEI GHT, LOWCOST, LONG II FE STRUCTURE, • CONVENTIO~L MATERIALS AND FABRICATION • LONG RANGE, HYPERSONIC CRUISE AIRCRAFT

Figure 6.- Flight test requirements for empennage structures.

SUBSON Ie COMBUSTION INTEGRAL ROCKET- ~JET RADOME

~'" ~ §"' ~ """',"0'"" ~ PROPULSION

NOSE TIP EROSION ANa " ~-~;J,t TP/STRUCTURES ~

~" ~ STRAKEIFIN ~ ~

LASER/PHOTO-OPTICS STORES SEPARATION '

Figure 7.- Other candidate flight demonstrations.

I lAUNCH \'lEIGHT, , PROPElLANTS, I BURNOUT WEIGH1, • MAX, MAC H NO ..

lDOS~lb 16118 Ib !lB1S Ib

S.06

~ ~-2J.Oft----! Figure 8.- Original X-24C concept.

LAUNCH WEIGHT. PROPEllANTS. BURNOUT WEIGHT, MAX. MACH NO.

WITH SJ

42394 Ib 27464 Ib 14930 Ib

6.1

NO CRUISE WITH SlI! M • 6

Figure 9.- First vehicle modification.

LAUNCH WEIGHT. PROPELLANTS. BURNOUT I'![IGHT. MAX. MAC H NO.

WITH SJ CRUISE AT M· 6

WITH SJ

49541 Ib 34000 Ib 1554llb

6.8

17 sec

NO FUSELAGE PAYLOAD SECTION

Figure 10.- Second vehicle modification.

Figure 11.- Payload bay-propellant tank trade •

LR-8l lAGENA) UDMHIIRFNA

RL-IO ICENTAUR) LOX/LH

2

Of--') ~','

".J

LR-9l ITlTAN) LOXIRP

UPRATED LR-99 IX-15) lOXINH

3

LR-105 (ATLAS) LOX/RP

SELECTION CRITERIA: COSl' PERFORMANCE ENVIRONMENTAL IMPACT

Figure 12.- Candidate rocket propulsion systems.

BASELINE UPRATEl' XLR-99 (X-15)

(WITH EXTENDED NOZZLEI

THRUST - 61 000 Ib

ISP-286sec

PROPELLANT DENS ITY -54.0 Iblft3

AVAILABLE - GFE

THROTIABLE

MAN-RATED

BACK-UP LR-I~5 (ATLAS)

THRUST - 8~1)()() Ib

ISP-3I)5sec

PROPELLANT l'ENS ITY -60.2 Iblft3

AVAILABLE - GFE

MAN-RATEl'

"IN PROl'UCTIOW

REOl'. MODS RESTART HOR IZONTAL START THROnLE

Figure 13.- Recommended rocket propulsion systems.

LAUNCH WEIGHT. PROPELLANTS. BURNOUT WEIGHT. MAX. MAC H NO.

\'11TH SJ CRUISE@ MACH 6

WITH SJ

554371b 39300 Ib 161371b

7.4

•

PAYLOAD BAY

~14.211~ "----- 48.511 -----

Figure 14.- Third vehicle modification.

~ DIRECT BOND INSULATOR/ABLATOR

<> NICKEL ALLOY HOT STRUCTURl

~ ABLATOR OR RSI ON CARRIER PANELS

~"'" ~"""" ALUM INUM OR LDC KALLOY

~'<rn","", LDCKALLOY

Figure 15.- Structures/TPS trades.

INSULATOR/ABLATOR ON ALUMINUM STRUCTURE

HEAT SINK STRUCTURE LOCKALLOY

TWO POTENTIAL TPS SOLUTIONS FURTHER TPS VEVELOP~!ENT REQUIRED

INITIAL COST RISK OPERATIONAL CONSIVERATIONS EXPER I MENT CONS I VERA TI ONS

Figure 16.- Recommended structures/TPS.

WING-BODY CONFIGURATION

B -52 LAUNC HEO

MAX, USE Of GH PROPULSION GEAR SUBSYSITMS

~~-'~"---L: ~~~c ~~,-. C=~I i.. 48.5 It !

• 24,2ft~

Figure 17,- General arrangement of USAF-NASA High-Speed Research Airplane (X-24C) •

Figure 18.- Inboard profile.

040 sec ROCKET CRUI SE AT MACH 6 • VEHI CLE TECHNOLOGY 01000 psI DYNAMI C PRESSURE

o AERODYNAMICS AND HANDLING QUALITIES

o IR AND RADAR SIGNATURE

o INTEGRATED PROPULSION SYSTEMS

• REPLACEABLE PAYLOAD SECTION, fiNS, AND STRAKES

o HOT, INSULATED OR ACTIVELY COOLED STRUCTURES

o HYDROGEN fUEl SYSTEMS

o WEAPONS DELI VERY

o AVIONICS AND OPTICAL EXPERIMENTS

RAMJETS

SCRAMJETS

INlETS

Figure 19.- Proposed research vehicle meets USAF and NASA flight research goals.

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X-24C Aerodynamic and Aerothermodynamic Design

The design program plan is arranged in a pattern of increasing model

complexity and cost and increasing eh~erimental detail and accuracy.

All tests can be run at AEDC. These facilities result in a model size - .

which can provide" the degree of detail and data for design purposes.

The plan provides for final vehicle configuration definition in January 1977.

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1. Quick Model: An existing wind tunnel model of the Martin X-24C-9 will

be modified to incorporate the significant shape changes which are now

considered necessary for the X-24C. The model is designed for force and

moment test, but because it is made of reinforced epoxy it will be useful

for limited heat transfer and oilflow studies. The X-24C-9 model has been

run in tunnels 4T and As so there will be data available to show the effects

of the latest shape changes. The fabrication of this model began in . --

January 1975.

2. Quick Tests: The latest shape i5 ..... a product of analytical study only.

Experimental data are needed to verify the configuration before large amounts

of money are spent on the next phase of model building and testing. A short

force and moment test. in 4T and A will provide these data. The duration will

be about 16 hours in each tunnel. An oil flow and heat transfer tests will

also be run in tunnel Bf requiring approximately 16 hours. /

3. Aero and Thermo Models: This model building effort is a contract

admLnistered by AFFDL. The purchase request has gone out for bids and the

successful bids-have been identified. Final selection of the contractor is

expected in late January or early February 1975. The contract start date will

be :'farch 1975 so some data from the quick tests are available for test

configuration definition. This effort will be funded in FY75 & 76.

The models will be about 27 inches long. A forcl;: & moment: model,.

pressure/heat transfer model, and thin-skin heating model will be built. All

control surfaces will be deflectable, using fixed-setting parts at regular

angles.

·;1

4. Aero and Thermo. Tests: A moderately detailed set of wind tunnel tests

will define the aerodynami.c characteristics of the vehicle, the heating

diStribution and location of high heating regians~ and pressure loading

for structural design. The data will be sufficient to begin full scale vehicle

design and flight performance explorations.

4T 80 hours; force & moment tests Moo !!II .4,.6~ .8, .9 9 .95, 1.0, 1.1, 1.3 Re·variation - with grit Control ~eflections. vertical surface rqmts.

4Q h~urs; pressure Moo" .49 .6, .St .9, 1.0, 1.3 Hinge moments h;,pm Cp integration

A 60 hours; force & m~ent tests Moo ... 1.5, 2.0, 2.5, 3.0~ 4.0, 5.0 I '.0 I 8~O Re·.variation - with .grit Control deflections, vertical surface rqmts •

. Oil flow~& shadowgraph.

40 hours; . pressure Moo =::. 1. 5" 2 • 0, 2. 5 , 3.0, 4.0, 5.0/ 6, f} j f3 I D IU.:nge Moments

24 hours of heat transfer data at Mach 5.0

B . 80 hours; pressure & heat transfer M.6·~ S oi.l. flows

5. Laniley Model Construction: Models for preliminary heating could be

fabric·ated in the Langley shops. These models would be of smaller scale and

used primarily for oil flow and .heat sensitive paint tests.

6. NASA Lan&ley Model Tests:

Unitary Plan 80 hrs paint and oil flow Mach 4

/

7. Large Scale X ... 24c Flow Field Models: Following configuration definition~

additional experimental work will be required using a model that will

incorporate more details of the full scale vehicle. A large model also allows

higher Reynolds number tests. The model would be about 40 to 50 inches long,

instrumented for pressure and heat transfer studies. It may also contain ..

remote drive mechanisms for the control surfaces. Associated instrumentation

would be a separately mounted flow field (pitot) probe and a porous plate

for study of ablation gas addition to the flow on the lower surface.

8. Large Scale Flow Field Tests: These tests will be run in tunnels A and B,

and 16 S if required. The tests will concentrate on flow field details which

must be accurately measured prior to final design of structure and equipment.

The flow field details include

a. Control surface hinge heating due to gaps and separation

b. Reynolds number scaling

c. Canopy heating

d. High speed drag brake flow interactions

e. Rocket exhaust effects

f. Local flows probing in region of experiment, includ:Lng scramjet inlet.

g. Simulation of scramjet flows with engine operating

h. Outgassing effects on experiments

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Proposed NASA TN-D

Flight-Determined Derivatives

of the X-24B Research Airplane

by Alex G. Sim

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REFERENCES

1. Sim, Alex G.: Fiight-Determined' Stability and Control Characteristics of ,the }12-F3 Lifting Body Vehicle. NASA TN D-7511, 1973

2.

3.

4.

Mecht1ey, E. A.: The International System of Units- Physical Con-", , stants and Conversion' Factors" (Revised). 'NASA SP-7012, 1969

Py1e,Jon S.; "an~ S~anson,_ Robe:-t 'R.: Li:t an~ ~rag ~arac"teristics ;" f) of the m-F2 L~ft~ng Body Dur~g, Subson~c Ghd~g Fll_ght'. Olev--e {jf>O'M... \.tto NASA TM X-1431, 1967 , " " ' '

Wolowicz, Chester R.; and Yancey, Roxanah B.:" Experimental Determi':'" , " nation .ofAirplane Mass and Inertial Characteristics • NASA TR

R-433, 1974

5. l.faine, Richard E.; anci Iliff, KennethW.: A Fortran Program for Determining Aircraft Stability and Control Derivatives from

. Flight Data. NASA-FRC H-856, 1975

6. Iliff, Kenneth W.; and Taylor, Lawrence W., Jr.: Determination of Stabi1ity.Derivatives From Flight Data Using a Newton-Raphson Minimization Technique. NASA TN D-6579, 1972

7. 'Norris, Richard B.; Sotomayer, Capt. William A.; and Fehl, John E.: , Parametric Study of an 8% Scale Model of the X-24B in the Landing

Configuration. AFFDL-TM-73-21-FXS, 1973

8. DeKuyper, R. E.: Transonic Wind Tunnel Tests on a .08 Scale Model of the FDL-8X Lifting Body. CAL No. AA-4024-W-2, 1971

9. Jenke, Leroy M.: Wind Tunnel Tests of an FDL-8X Double-Delta Space­craft Model at Mach Numbers from 1.5 to 8.0. AEDC-TR-71-218, 1971

10. Lindsay, Earl E.: Aerodynamics Characteristics of the AFFDL X-24B Configuration at Mach Numbers from 1.5 to 5.0. AEDC-TR-74-87, 1974

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//A..rA TN /)-7S-1l /'173 j

I ! I

"i

.~ l

i L

". f::"­,.

, . , L-_ .... _ ; .

~

36

NAJA TN j)-7.>/Fc

(M~-i:3)

Figure 3. Sign convention and control surface location.

~ )

lyt1.-FJ TABLE 4.- PARAMETER RESOLUTION AND ACCURACY

/l

Parameter

q, hN/m2 (lb/ft2)

o , deg a

o , deg r

o , deg u .....

p. deg/sec

r, deg/sec

{3, deg:'

cp, deg

r, deg/sec2

a,g y

q. deg/sec

0:, deg

e, deg

• 2 q, deg/sec

a , g n

a , g x

Resolution

0.670 (l.40)

0.111

0.09·7

0.111

0.0594

0.157

0.050

0.040

0.380

0.829

0.380

0.00539

... ----. ';"YJ·

. 0 .168 (otA)'5~ -~-- ~ .

0.0607

0.187

0.349

0.0174

0.00870

Accuracy

1.57 (3.29)

0.675

0.380

0.462

0.675

0.462

O~830

0.550

0.220

2.48 .'

0.0164

0.550

0.43

1.24

0.0328

0.082

"

31

• >

J.t~-) .­~

.. ""--

a, deg

0a' deg

. {) , deg r.

p, deg/sec

l:r r, deg/sec

8

0 13, deg

-8

4

cp,deg

-4

20r. ... o \" ~ "20: VVSJS· I

a , 9 Y - --

O. 12 .. 3 4 Time, sec

4

0

-4

1

0

20

0

-20

.1

-.1 .0 1 2

f/A-JA-- TN lJ-7Ylf .

(/'/12- F3j .. .

.

3 4 Time, sec

Figure 6. Typical doublet control input maneuver.

39

;

.~ J

~ .)

40

a. deg IT 0

16

8

Oaf deg 0

-8

-16

4

Or' deg 0

-4

20

Q ~

P. deg/sec

-20

-40 0 1 2 3

Time, sec

4

r,deg!sec 0

-::-4

1

0

~, deg ~1

-2

!p, deg

.1

ay' 9 0

.... 1 4 0 1

;YA.f A- IN f) -7.>11 . (;/l2 -;:3)

'2 3 4'

Time, sec

Figure 7. Typical pilot-induced oscillatory aileron control input maneuver.

. :":-.

.... ~ '-: .. "

..

48

-1 C1 ' deg

13

-I Cy , deg 13

o Flight Wind tunnel

-_012,

. -.008

o

-.004

o ~ ______ J-______ ~ ______ ~~~ ____ ~ __ ~~-J

.012

.008

--: r I 0 0 <e> (l!) <,t!!) . 0

00 0 00 0

0

I -4 0 4 8 12 16

a, deg

(a) Cl ' C Cy

_ {3 n/3 ./3

Figure 11. Comparison of lateral-directIonal derivatives obtained from flight data with wind-tunnel predictions for a Machnurnber Qf 0.7.

Z~~tf1'ffffEii!'t~Wit1£ff@5~BiiiW~1i%fW1i'4~~W~¥~F~~'rm.~~ -~. '. '.' . ...... '. '. '. .' .... . . ··HF~

," ~.., . . '. .' '. '. #.4.JArN 1/-7~-;t ··W~ . , ~~

o ~ ~th%-1 r~

J ~ ~b;

h I· .. ·, "',

~ ..... . J'

i ! i i

·1

it' f~·/ I

o Flight --- Wind tunnel

o 0.8 0 0 Q o o

.~ ..

"-I ;~[<> . . Cy .' ,neg '.'0 ~---n-+---_...o---O-<>-----n--O-"";-'-"""--':""-

6 o~o , ··a

-:'.002 L-_-:--~L-__ -..J. ___ ---L--""":"":"---:---l-.-':-_--':---l

-4 o a, deg

Figure 11. Continued.

49

"I,.~ IflJ IH I ~r~ II! ~ fi1.~

J.fl . ,. Ikl-IlltJ ,tlt~

~,,~11H ""- ~fr~ . 'f"F ~

.{!f ~

~p ~

~ll

Jf:r:~ ~ J:~ J:~' ~ ~ .. , ,

,!- :;~,

f6i~'" w;

:: >

t--;

§.,. -) ./ 50

.. 001 o cf

(')

Flight --- Wind tunnel

Flag danotes b = roO sb

)lA-SA /AlP-7)/J

I M2 -f:3J

Slash denotes outboard fin . jettisontubes on

~.OO4~----~~~~--~--~~~~~_-L----__ ~ -4 4 12 16

a, deg

Figure 11. Continued.

I I I·

t I

J I·

. t

f i -

! ! f I

I j I

j t

') !

-1 Ct ' rad '

p

-1 rad

.4

", -':

o

o o

· ... ~" :~-:~o

o

C1 , C

p n

p

Figure 11.

•• -- •• -------.•••• - ________ ....... ~_ .... ____ ......... ~_ 0:: 1 ..... \!CIk_· .. "'!;5.:,.~

o

aoce 0,

o ,0

1-,', l~

o 8 .;: .... , o

c·' C 1 ' n r r

Concluded ~'

Flight Preflight estimate

:,00,· , ,

I

'#45.04 )'N P-7571 .lh:~ ( PH-FJ)~;~\:

~.;.~.- ~

,.,:

t:. . ~ ~ . ~%.~ .. i.:t...": . i!:~' ~ ; "', t~-' F!.~

'it F t 1!:-; ;z!:;

Iwi !", "'-If,-

", :'~"'l

~~!,

L'".:."

52

-.012

de~:l •• ·. -.-

,'-.0.2 r' . l' 0. .

.. ; .. -4 . 0.

(a) C1 ' {3

'::-,..:.,.

-'. -- ... -",

C , n{3

a, deg

o Flight Wind tunnel

o

o OJ ·'0'" o

f 12

;VA.fA rN[)-7rlf / /4 2.-F3}

. ·16

Figure 12. Comparison of lateral-directional derivatives obtained from ,flight data with wind-tunnel predictions for a Mach number of 0.8..

I ! t f

j I

I

- ':',~, 0 -:~:-~->~-L:;,.·::~:<:"-~;·~i;r:~':~-.~ --_ .. " ~ . -;' .~.

'U/;'p;;bl;YA~b5-'

"."., ," ..... :.

4 PM'

-,

"-~ -q- -I ('0 ... fl • x

o (J ()

f})

Q () 0 CJ

~ r-

\J

".-_. > "0 • ~

~---J -. --.-. 0 () '" ~0-

0 I o. \8 t.;j 0

o 0.6

1 --, f ,­,_ !-~, { - U' '..:1 . . . ,,,- ,

0 .... 0° o '0' . -

·:.r

0.7

---~-- ....... --;

'-

0.8

/Vl

---- - "_._--­~ -"--

,- . -'--.---~'-----' .

'~. _--- ---7.' ".-'

.~

0·'

~

~ I

""' . v {)

(; -0 , .. 'Jj

o

. 0 o I

-- -~-~-------- --:-- -:--'-

.:.---- --

---,-

_':"_------ .

"""

".--- -----~----- ------~-~----.-, -.' ....... ~ .. _-------'-------:-.

/~\ I ' , )

'L ..,) o 4 Jb

J

e.G. -= ,6"S""Z _

..-.../ ,~ i

tw··

f"', , \

. t~OTE:' ALL CONTROL SURFACES

Definition and Sign Convention

19

)" 24C- /2)(-1

/3/ /

_Body

W//;9 -... -.----~-,,~,,~-.-- ... -~.- ,,--,- -_ .. -... ----.. --.. -.------~. -.-.-------------------- . -- ~ .. - .. -- .. ------ ---

f

C-e.-ItJe. r: _){-I-)/ __ CA /

----_~_____ ~~2-7-_-------€/--d,o#----.E/L7~\S'-~----- C)~o_e.-:~ ______ ~ __ -' ______________ . _______ ~-'-....::.....----

~_-~~=_~~-=~~-tsl.~~~:~:-_-:- S')~d~:_ ..&~.-~=~=~~--~-/a D-~~~~ --Z:n~~=-~------~ -'.' --- ~~_~~-__ ~-- _ --~~~~~-----------,_. __ .. ______________ £~ ___ -J-----------Po-dy _____ £/0_ f:).~_ _ _____ . ___________ . __ . ________ . __ . ________________ _

T"--- , / 7 . ----~----.---- -.~-- .

. _- -- ----.--.-------~--.. ------- -~ /, ----. - ... - .. -._----------------- ---.-.-

.---- . Je.., 7 __ . __ .J.-, e. VOJ1 _____ . ___ -. - -. ..--

--- ------_. -----'---- --.- -----------,-... _---------- --------'-_.,--_ .

/ \

j)--- -v'- ,.---.-.-... - ...• _- "'-"---.-. j

S_V..6t/t~/i' £et 0---,D ~+.."";Rodj~7f i I ' ,A

- ....... s.o· /

V i;/t'_',I. 1-17 --/--

I! It II /-

/

I I

!!'C I /)8 /

-- .-._ .. _ .. _. -

C-e!' Le;.. ! J ! ;' V e..;-- t ( C'.._ a I

-'-- . /,>f./ , ;.' /

I

r..

I

\. )

.4l> -'1

)

. ,'. 15/W-' ~ . /vi-AJ!:>

INTRODUCTION

Fu tureuse of ai rbreathingai rcraft requires that the hypersonic speed

regime beil1yestigatedusing a hydrogen fueled engine to power the necessary . '. '.: . . . . .

. research aircraft. (ref. 1). As poi nted out in reference 1, J iqui d hydrogen can ~; .

illsObeused to cool the engine and critical parts of the airframe, thus . ".

'. alleviating;,sorne of the major heating problems associated with hypersonic flight.

.Anotheradv~ntage is that hydrogen fueled aircraft will help alleviate the oil

ShO;;ge01?> . Hypersonic vehicles will have an important role in future civilian trans- .

port.ation systems by r-educi ng fl i ght times, therefore moving more. passe ngers.

The military will benefit~;Y reduced intercept times .• $.Tii5!!§ =;bil" ~el1efas

These reasons i ndi cate the ne~d for a hypersoni c research ai rcraft as a step.

towards. rout-'nE: hypersonic flight.

Prev~0Us st~.dies (ref. 2-5) by McDonnell Aircraft Company identified the

/2quired ; ._':;2arch for future hypersonic cruise aircraft. These studies re-

CO: .. :',d1dec :r::..: ground facilities and flight vehicles be utilized to develop

lars;: Jod"·,,,:; aircraft would be necessary in order to transport the required

liqui: .. hyc.rogen fuel.' -:-he flight research vehicle would be programmed to be

6- '[' l;::unched, rocket boost to cruise speed, cruise with power supplied by a

.- '·.;2n fueled scramjet engine. After cruise the aircraft would be flown to

_ c~less landing.

-:-;,ei'l:!-:ore. to insure a safe flight. from air launch to deadstick landing,

".:"

.:2 "crojynJ.:nics of the complete speed range must be known. Reference 6 reported

'-':. i.2rc-!'inamic characteristics of an all-body supersonic research aircraft

::.:;·:.;:;ur,:.ion at Mach 6. Reference 7 docummented ~he low-speed aerodynamics

t>t~ ~ l~ • jwv--2 •.

') /

·ofa~65" ~weptdeltawing hypersonic aircraft, while reference 8 reported t~e """ "," . . ..... :

lO\'J";sPeeda~rodYnamits of a 70° swept delta wi ng ai rcraft .

. .......... The, purpose of the present report is to present the results of an inves-

··ttgati:§ndn~othestabilityand ~ontrol characteristics of a 70 0 swept delta

res~'archaircraft at Mach 0.2 fO.r ,1 Reynolds number range, based '6 6

X 1 0 to 21. 8 x 10 •

).'

'i,

" Ii,

.i .

"

,;i : , :

." "

'.\ J

;:,

. __ .-

... _If? __ _

......... #·;1.'-SI.i.i."M·.:.~· .. I.·:A.~.··.{j:.'-·. :··A.····~6.'.-.~./ .. A·/.·.~'.W ... ·" "\' ....... . . z7J" .{I(] . ..... . ... ";' /V-t. C r I /\. -::::/lJO? /t;.-t ."-? )... .. ·,.·.J-.O.t. .. :· ..... ....:::... .. C-I:- - ... "' ~. 'j4Y3"· . .. . .. . .. .... . :/ . .. Ct ---.It. - ..

.... _ .. ' ....... (211."2-

. #,~",.J fl_-~.e_P0J7_' . ,-_ ·_· ... 1.···,::..· ..... j .. ~ .... ; 'j"..~ . . .' I'" /;.. ...... }\ .rd ll

tJl~l ,: ·~r :_.~.JJ"Y CV:vrl~V 1 .. .2:..)<;.T~-)':yOrl~-;l)2:-t·3vr~'---I.B9/=ih'f~····-

.. t1.'.¥ .. ~ .... ! iJ/)--" .. .; I.r-j ·/·:lcf+ S->7;j -- ..... .~~ ~ ••••••••• ~n ••••••• ~ d ••••••.•• ~ .• tq .. ~_~,.-.7.£(/ ... P-

&/l ~S(~~~· -==(;.~7> t- li,l,C)CJ.9r l

,'-. 7·)'1/.;' 7u~l;q,,~ ~ . '. .. ... ... . .. :.z ..... .. ( .. D.) ";:.. 0' ~ . z. 17 ...

A 0 Jut f)ll~= (~~~Lt_~ -1~J:~) /0.17 . -- 7.)f 1~7 '1." ~ ~ .. . ... . ..... . . .... ~2- . .. ... ( 0 ') ....? (P.,7 r .. .

/v'L c-~/~ .. ;,t1&b £./1.> >4-~;- ~IJ4(k·+ ILj.]..)" )/(W -.1.f?1 ::-:r't. [)l7i ~y - . - l b ) :::- ?t'l. r9r"

/t1cu/70-s,4\::: % ..Ill- '{'If)'" ::/,n~7·;) (A,,~ .. t iP;L i. Y9 .- .. . .. or; rtj( -

~~'-' .- -- _ .. _. - -- •. _.- .-~. . .•••• _, ....•.. _~" __ ~ .••. "-._0.,.",_--. . __ .... _ ... __ • _ _.. _" •...... _, ___ .. ___ .¥ •• _,~_._ •• ,.... .._< __ .••... ". ___ •

J ~ fl:C§~nTA0~y,X-,2t1-C~ !2TdSdA5Ad~I1"dJY-e_~-t:. _~ __ X~lili __ 1-<---F

f-E1?L, / _______ _ _______ _ ___ __ __________________ ______________ _

- X 2') ~ _ ... __ ._ .. __ .. ~- .. - .. _ .....•.. _. +----- . - .. ,-~.-.- .... 7-._. __ .. -- " .. -.. -"------.-- "_." "-, .. ~ .. - .... ---.--.-'~-~ ... --.~-.-.--." ... _0 •• _ •• A_"" ___ .-_ -_ ••• _ •••• ~.-.-.~-•••• - •••••• _ •• __ • ___ ••• _. __ .~_. ___ ........ __ ••• __ •• _. ___ • _____ ._, ___ ._ •••• " __ •• ..,. _._ ••• _0 •• _. _____ • __ ._ •• _ •••••••••• _ ••

~ m-t:nfl~vtS/cI-e_ -- ;::;lL~d--!)i/;~fI?:?ll __ ~s:_dE_/)ictC!.t-e_~_ ___ _

.. _ .... __ ..•.• =:_DJ~~,~~~s!:;.~~2·~~~~ .. ··~~~L~t~~~~§=~=~~:........._.~.~.-•••• _ _______ _______ = _______ ~'. _______ W#£A=_£Z ___ . ______________ ~ ____________ :_' ____________ ~'____________________ ___ .. _________ . . -- ----------- .. ---- .. "---- -. ___ WItEA __ =_ Q.L __________ - .. --~/-----J2nlt_---------------------.---------.-------- _.___ _______________ _

-.... __ .~.- -_ ....... _--- ... _-.--_ .. _--_._---._--_ .. _ ..... _--...• __ ... __ ....... _-_ .. _--------- .-- ......... -.- .. -.--.... - .. ----... ----- ... - .. ~ .. -.-.. ---

.- .. -----.- .... - -_ .. --+ --~-.-... _.- -.----_ .. -._- - . - ...... _._._. __ .. - '-·-·····_· _____ ·_·· ___ .. -' .. _A.~ ... __ ._. ___ ........... _ ..... ~ __ .~ ... _. " ... ___ ".

-- - - -- -- - --- -- ._" - ••••• .- •• --. _.- - <" .- .- ••• -.~ •• '-- ." ... ~-••• _-- -- --.---•• ---.-.----- •• ---.-------~ •• " .--~ •• _.-•••• - -. ~-.. -- •• , •• - --.-••• -.,,-••••• -

_ ••••• _ •••• - •••• - ••••• -.- • __ .-._-_ •• -. - •••• _ ••• _ ••• ,., ___ , __ •• __ "' ._,,_ •• _. __ ••• 'k_ • __ •• -. _____ ,~, ___ • _______ •• _ ...... , •• __ .. _. __ • ___ .... __ • _, _ -- .- - -- " •• " -- '. -_ -_... " • _. __ •••• _ ........ ____ ••••• _ "' ___ ••• _ .~ __ " __ • ____ ~'_.'. ~ _' _____ "' __ ' __ '.

..... )

. _---- -- .. _.-.. ,," ---- -_ ... -- "_. ------

. _ .. ,. - ,. - ......... _.- .... _-_ .. _-_ .. _ .. ----_._.- -. - - ---,,-_. -_.". .

- ··f'tJ21~£tJ)~wt:tz;z:rw:~:s;rf~· .• . -· •• ·· .• __ ••. --.· •• jjQ~==~~_~lk;p~~;J/1·.-~:._ ...

,)

)

(

,"' N

,~,

PAYLOA08AY

SECTION B-B

,

~~~ i . -----tEg:' I

I·· ....... ,

: LO>' \ -,- '. \ '~ ":~'},,, : .)g).-PAYLDAD8AY / --- 't-r:~ \

=, - . ..) , , SECTIQN A-A

r' INSTRUM~NT~ \ I,

rWD 1!o!12 ATTACH PT.

, ST ... \ IZ2

~I----------~---

,v

1'31

~ I~rua~, ----.;.:--rr: - ..

, STA STA 3~O 344

~~~I'lOMo(I"'t

UPPER FLAP (2)

TURBINE EXHAUST

2~F"T = ].~~

13'1

ROCKET ENGINE YLR'99 RMI

LOX DUMP

LOWER FLAP

UPPER RUDDER (20· +L, -R) --

I LOWER RUDDER (10" OUTBD ONLY)

!

REFERENCE (PLANFORM) AREA LAUNCH WEIGHT PROPELLANTS

OXIDIZER 8,642 FUEL 6,913

H202 585 (APU) (176)

OVERALL SPECIFIC IMPULSE BURNOUT (LANDING) WEIGHT. VACUUM THRUSt MAXIMUM MACH NO.

(WITH MAX II = 400 psf)

'''./

• INCLUDES 723·lb GROWTH ALLOWANCE

UPPER FLAP (2) )7 (·40· UP, +20· ON; .

I .

~

616.4 ft2 30,053 tb 16,140lb

269 sec 13,901lb 58,6001b 6.06

-----~'-J -,-1-:1" '~>I/: -~---:.:l:.:- "-I J::s It -~ .! .!. I " WL.O>--- Pi i"'¥-

1/ i A ' 7 , APUW.J' _: ';-"------110----

~~~ PAYLOA,o BAV .,

-!'"4!)' r~E MAI~GEAR{MCOtF'lED) ,., 46.1 FT-----,- -"--"--" ;;;:..,...0. .... --.. ----..... _, __ ,_. __ _ 4NGINE SmVICE P~N<L

FIGURE 2 BASELINE CONFIGURATION'GENERAL ARRANGEMENT

17 and 18

x. -)., t) -C.... q

... ,

, ,

~.

station xll

'V'

r .. ~.: I .-·IF'~== .. ' ~- . --';1-0)---I -'T I I __

Station xl! , 0 .124 .233. • 374 .499.599 • 710 • 797

' I I· 6~6 I I

,I· I '

, (a) Completa inodel.

.909 1.00

I

Figure 2,- Details of wind-tunnel mociel. All dimensions are normalized by the fuselage length L I . . :,

IIYFAC,

~/ I .

"

50f

.395

ct 'lI ", Ii') -) ..... ~

~

.....

fj

....... ,~

I ~ 1,...1

~ ~ -:;;

'" '"j..

2 00 ~ .....

'C\( $ () ~ ~ It ~ , 0 0

, ~

r- "::f.. ~<t ('(

)

J

~'----'.

, .

xll .. .124

-;-.233

.374

.500

.599

.710

.797

~-.909

~--.946

,----1.000

'\..J ,

<[I::::J-l--I -, -

.Lr· c --I -zzt::::-

!~ c/3 1101. c/3· .. I. c/3 j .05 c

o .124 .233 .374

t

i&

STMION xll' .500 .599 .710 .797

I I I I c.g~

I

.~ RETRAC1tDSC~ FIGURE 2. - DETAILS OF WIND TUNNEL MODEL; DIMENSIONS SHOWN IN PERCENT BODY

LENGTH. l. i

VJI;)1 r-fJ J II )I FA c..

..... --+--~

"":

'I

,/

i, ,

"'( ~ ~-

'" . $ 10) C,> ..f"{ ) -'"'" ~ f")

/ ~l4 ~ ~

-t....

~ ~ Q:)

{'(

0 0

-.0 \n ~ <:) t' \'( --.. ~ )C

H \\j -

~ II

1 ~ ~ -...l

<t C\( Q::

" .~ ~ I ~~ .

U·

" " ~ ') '6 ./ ""\0

r-.t . '-J ....... ~

~. ~ 't;.

q)

~ ~ ~

1 I '0

'0 \..n ~ tt1 C'-! -... ()

) ~\c -'/

L

Station 1.0 _Q .901_tlj_

.833_ffi-· .765

Airfoil section'at wing tip

.026

'---- Dashed lines show negative cambered airfoil W2f

\J

c::==== --~

Vertical tip fin ai tfoi I section

_.---___ ;-_\ -'-==" Station - ~~ .219

~ ('

L-.b:=-- . c. g. I .~, ., '0 0.0

.208 Station

"(al Baseline configuration. All dimensions have been normalized by the body length (i.~ 50.8 em) .

. . :. ... , .-.. ", ... , Figure 4. - Model general dimensions.

. . \

WJ.lRA -5-2-

~ __ station , .?19

'!;t~ff~~j_~WWtf"tf1X wren., • ...," jZ' Ur" .. 1 __ st .~ .............................. '"'"'-~_.~. _,-,~ .... __ . .J..~""_'."'~"';'.""';"""",,_._~":""'_''''''''

h. V)

~ "'.l f Cl.... ,--, r- h ) ,

Q... ~ f- " ~' -......! co

~ ~ 0 I!I • ~

~ ~ 0 ~ ~

• ~ II

~ --..l

<l.

)

"

'':'" ~ \

~. Cl:::

~ ::t: ~

J

~ . . ~ 0--..

~~ ~

..~ ~ 4.t. <)

~ - In - 1('-) -..:lI. C'\.J x i ~ "'h.

>: " ~ I N .....,. I' 'c

.3 ...;, e 19 1 ~ .......... i

<:> ......... ....( if i~ ....... x <:I ~ I I

I~ ~M I

·0 o· I f"{. ! 1

'-t) 1<"{

o--~ k • <:)

~ - ~ II x

I ~M <:) '" -"J. ~ t'{ "-l <l If

~ 0---.. . ......

~ ~ ......n '" •

I ~<6 h

~ ) I .....

t"( <"{ -..... ~ I ~ ..

00 ~ ~

)

~

-.

.~.

8-52 LAUNCH AIRCRAFT CONSTRAINTS " .-.

9.0 MIM. CLEARANCE

18.0 NOM. CLEARANCE

CPNSTRAINTS

WING SPAN IV 23 ft FUS HEIGHT IV 9 ft . GROSS WEI GHT "-' 110 000 Ib

co. '~®30: . 1 _ --~- . t­

COMPRESSED GEAR GROUND liNE

STATIC GROUND LINE

d\-. .

~.

I ---

I __ i I I

J

1

-~,~ -- ' -r" I --+--

; ;.

~i

..,;

) ./

I .1

/ /

I­L1.J --, ~ « 0::: u V')

::cl U l

~I ~I 0::: i

t

I \

\

\ I

\ !

\ \ ,

)

1.0

~. -

. ")_."

.. -1

- r.- --- _. _ .. -".-".-'--.. "'--.--~ . __ ."-"- --"-. __ .

. , -;

-,---~ -'-!.-- ~ - -~ --.. - ".- . -

." - ) - ~ < - ->- -. •• - • ,.. • -.~ ___ +~_o __ ~~.:-+:~.:~-~----- .. ~ --(_:.:.~~: -i

" ,

----"_. .. - -----.-._-: .. -- .. -- - - -B---- -.... ---. . g I TPr -~ 7;/,/. ' -.----: -- -;

'-----------:---- --------~------- '-l---.-._' ------'_. -.---'"-.~-~- .. --~----.:.~-.. ---

o~----------------------------~--------o .4- .. 8 /.0

\ , \

\ I

)

.... ... N

1"<:----------'' -l'.! i . .~

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