The x-15 flight test instrumentation

T€E X-17 FLIGHT TEST INSTRUMEPJTATION By Kenneth C . Sanderson NASA Flight Research Center Edwards, Calif.

Third International Flight Test Instrumentation Symposium Buckinghamshire, England April 13-16, 1964

MICROFILM $ . . /a2 y

April 21, 1964 .

THE x-15 FLIGHT TEST INSTRUMENTATION By Kenneth C . Sanderson* INTRODUCTION I n 1934, t h e general requirements far a new research a i r p l a n e destined t o follow the X-1, the Douglas Skyrockets, and t h e X-2 were established by the National Advisory Committee f o r Aeronautics. The major g o a l s of t h e proposed f l i g h t research program w i t h the vehicle were t o explore aerodynamic heating problems, study s t a b i l i t y and control problems i n a region where aerodynamic f o r c e s are n e g l i g i b l e compared t o i n e r t i a forces, and explore physiological f a c t o r s a f f e c t i n g t h e p i l o t , such as weightlessness.

The configuration decided upon as best s u i t e d t o meet t h e program goals i s shown i n f i g u r e 1. This configuration--the X-l?--is capable of speeds g r e a t e r than 6,000 f t / s e c and a l t i t u d e s exceeding 250,000 f e e t .

The powerplant i s a rocket engine t h r o t t l e a b l e from 28,500 pounds t o 58,700 pounds o f t h r u s t w i t h a burning time of 80 t o 8 3 seconds a t f u l l thrust.

Liquid oxygen i s used as a n oxidizer f o r l i q u i d anhydrous ammonia.

The major components of the primary s t r u c t u r a l elements behind stagnation points were capable of w i t h - standing temperatures of approximately 1,200' F.

The performance envelope o f the a i r p l a n e i s shown i n f i g u r e 2. The wide range i n dynamic pressure from l e s s than 1 l b / s q f t t o approximately r 2,500 lb/sq f t , coupled with high speeds and high a l t i t u d e s , generated formi- dable problems f o r t h e instrumentation design engineer a s well a s t h e a i r c r a f t designer.

It i s Some t y p i c a l X-15 f l i g h t paths a r e depicted i n f i g u r e 3 .

important t o note that t h e X-15 would, during i t s f l i g h t program, range over *Assistant Chief, Data Systems Division H-339 -2 - distances as great as 400 miles, covering a three-state area in the western United States. Hence, a long, well-instrumented t,estrange away from heavily populated areas and major air-traffic lanes would be required.

PROGRAM INSTRUMENTATION PHILOSOPHY The basic instrumentation philosophy for the X-13 program was dictated X-15 were to successfully f'ulfill its primarily by two factors. First, if the mission of providing timely research data, it had to be built and instrumented quickly. Second, the instrumentation had to be accurate and reliable.

The philosophy adopted was as follows: (1) Onboard recording would be used all the data sensed on the aircraft to eliminate the as the means of recording risk of data loss and degradation inherent in radio-frequency telemetry links.

(2) Inasmuch as the X-15 would be carried aloft and launched from a B-52 aircraft, selected parameters, including engine, control system, hydraulic system, environmental control: electrical system, and pi1 ot physiological data, a would be telemetered and displayed to ground monitors in real time to insure safe launch and flight. (3) Continuous ground radar tracking would provide information necessary for ground control and would also provide a source of space position and trajectory information for research purposes. (4) The instrumentation system would have to be flexible to meet the changing require- ments of the flight-test and research engineers conducting experiments during the flight program.

(3) Finally, maximum use would be made of off-the-shelf instrumentation components and systems and existing facilities to f u r t h e r maximize reliability, minimize costs, and enable program schedules to be achieved.

-3- BASIC INSTRUMENTATION RFQUI-S A req2irement. for 1,000 to 1,100 measurements on the airplane was arrived at through ar, iterative process involving many groups interested in conducting Based on constraints imposed by the manufacturer of 8 0 0 pounds experiments.

of general instrumentation, 40 cubic feet of space, and 2 kw of power, a total of 800 recording channels and 90 telemetry channels was decided upon as the best compromise between the research and flight-test data requirements and the constraints. A ground test range of three stations capable of providing con- tinuous tracking, communications, and telemetry; a flight simulator for pilot training, flight planning and data analysis; and a digital computer f o r theo- retical computations and flight-data processing rounded out the basic instru- mentation requirements.

Design Measurements List The design-measurements list consisted of 1,050 measurements distributed as follows: Research

Skin and internal temperatures . . . . . 588

Strain.. . . . . . . . . . . . . . . . 64

Control positions . . . . . . . . . . . 28

Aerodynamic surface pressures . . . . . 1 3 6

Basic flight parameters

( a , f3, ps, 6, ir, C ) . . . . . . . . . . 22

Flight test (subsystem) . . . . . . . . . 212

Total . . . . 1,050

Most of the measurements, as would be expected, consist of structural temper- atures and aerodynamic surface pressures. Approximately 200 measurements were required during the flight demonstration to verify and test subsystem perform- ance. The design-measurements list was formulated on the philosophy that the -3- measurements would serve t o meet f l i g h t - t e s t requirements f o r design v e r i f i - c a t i o n and would a l s o provide data f o r research purposes.

AIRBORNE INSTRUMENTATION Instrumentation Locat ion Figure 4 i n d i c a t e s t h e l o c a t i o n s of surface-temperature and pressure- measuring i n s t m e n t a t i o n on t h e X-13. The wing measurements a r e concentrated i n t h e r i g h t w i n g t o provide s u f f i c i e n t coverage and s t i l l s t a y within t h e i n s t m e n t a t i o n space and weight l i m i t a t i o n s . Coverage on t h e nose a r e a and t h e upper v e r t i c a l t a i l i s heavy. There i s no instrumentation i n t h e lower half of t h e v e r t i c a l tail, s i n c e t h i s portion of t h e f i n i s dropped before landing.

Instrumentation equipment is carried i n compartments i n t h e nose, j u s t t o t h e r e a r of t h e p i l o t , a t t h e c e n t e r of g r a v i t y between t h e f u e l and oxidizer tanks, and i n t h e t a i l s e c t i o n ( f i g . 3 ) . The main instrumentation compartment i s behind the p i l o t . A l l instrumentation w i r i n g and tubing behind t h i s com- partment i s routed through tunnels running along each s i d e of t h e a i r p l a n e .

The main instrumentation compartment and t h e nose compartment a r e pressurized and temperature-controlled. The center-of-gravity compartment i s temperature- controlled, and t h e t a i l compartment i s insulated a g a i n s t high temperatures b u t not pressurized or temperature-regulated. This environmental c o n t r o l configuration w a s d i c t a t e d by t h e types o f instrumentation equipment i n s t a l l e d i n each compartment and t h e e x t e r n a l environmental conditions of each compart- ment. Individual instruments and equipment a r e shock-mounted o r hard-mounted, depending on t h e c h a r a c t e r i s t i c s of each u n i t and t h e v i b r a t i o n and shock conditions a t each l o c a t i o n , I n s t m e n t a t i o n equipment was designed and con- s t r u c t e d f o r hard mounting wherever practicable t o save weight and space.

-4-

-5 -

Instrumentation Design Considerations The design and construction of the X-15 provided the aircraft designer with the task of designing new structures using previously untried materials and new systems capable of providing control where inertia forces were predomi- nant. The task of the instrumentation engineers was equally, if not more, difficult. First, the instrumentation system had to be accurate and reliable, which meant, simply, that it had to be an in-being operational system at the start of the flight program. Secondly, it had to produce the required data in an environment where the sensors and wiring would be exposed to temperatures of the order of 1 , 2 0 0 ' F and atmospheric pressures as low as 0.03 lb/sq ft. Early in the instrumentation design phase, a number of problem areas were delineated that would necessitate development programs or would present difficult design tasks. It became obvious that the weight, volume, and power constraints placed on the instrumentation system would present a severe, if not insurmountable, problem. Measurement of angle of attack, angle of sideslip, dynamic and static pressure, velocity, altitude, and attitude would not be possible with available Structural temperatures and aerodynamic surface techniques and equipment.

pressures, which are of prime importance, would be difficult to measure to the required accuracy and precision.

The selection of an instrumentation system which would meet the basic X-12 requirements and philosophy and continue to do so for 5 to 10 years required careful consideration of many factors in addition to reliability, accuracy, weight, volume, and power. Of concern were the cost, the design, fabrication, and test lead times, the capabilities of facilities and required operating personnel, the difficillty and probability of s o l v i r i major prchlerns, and the data-processing effort and time required to present the data in useful form to the flight-test and research engineers.

-5 -

-6 -

System design.- The instrumentation system selected is shown in block- 6. The system, which relies on oscillographs and pre- diagram form in figure cision photographic recorders, was chosen f o r the following reasons: Most of the components were readily available from commercial sources or NASA stocks, which helped to keep costs within X-17 program funding levels. The lead times for designing, fabricating, and testing the components and systems were con- sistent with the requirement that the system be operationally ready at the beginning of the flight program. NASA instrumentation personnel were thoroughly familiar with the operating principles, service, and maintenance procedures and could be drawn from the NASA staff with a minimum of training, thus saving time and money. The difficulties of adapting, where necessary, the available tech- niques and equipment plus the probability of successfully accomplishing the required developments were again consistent with program schedules and costs.

The time associated with processing the data from an oscillographic system, especially where masses of data are involved, was, and still is, long compared to the automatic techniques which can be used with magnetic-tape systems. This element was carefully analyzed, and it was concluded that the estimated number of data points required per flight (15,000) would not create processing times that would be detrimental to the planned flight schedules. This, coupled with the fact that during the 1956-57 period a costly, time-consuming development program would have been required to obtain a fully automatic magnetic-tape system, made the oscillographic system the choice from the data-processing standpoint. From the reliability and accuracy standpoint, the oscillographic systems had been proved in flight.

A s shown in figure 6, aerodynamic surface pressures, linear accelerations along the aircraft body axes, and total pressure from the nose sensor are sensed and recorded on precision, NASA developed, electromechanlcal self- recording instruments. Outputs of angles of attack and sideslip from the nose

-7 -

sensor are recorded on precision NASA recorders employing servo-repeater systems to position a light source on moving film. The attitude-angle outputs from the integrated inertial flight data system are recorded in similar fashion. All other measurements are sensed with electrical transducers. Signals are col- lected at a central patch panel in the main payload compartment, routed to appropriate signal conditioners, and then to the oscillographs and the telemetry set. The oscillographs are NASA developed, 36-channel units. The telemetry set consists of a 90-channel pulse-duration-modulation system that uses a low- level electronic commutator and an FM-FM system for telemetering pilot physio- logical data. A low-level mechanical commutator was used early in the flight program. Where required, parallel outputs from instrumentation signal condi- tioners are sent to the pilot's display instruments. Recording speeds can be varied from 1/4 in./sec to 4 in./sec, which gives recording times ranging from 5 6 minutes to 3.5 minutes using 70-foot film magazine loads. A blue-sensitive polyester-based thin film with the trade name of Cronar is used. Sixteen- millimeter motion-picture cameras (not shown in fig. 6) photograph portions of the pilot's panel and the wings and empennage during flight.

Air-data and attitude-angle measurement.- Measurement of angles of attack and sideslip and the pressures used to obtain airspeed and Mach number in the speed and altitude regime in which the X-15 operates is limited by the ability of any instrument to withstand the high stagnation temperatures and the low pressures and lags at high altitudes. Thus, conventional methods of measuring angle of attack and angle of sideslip using nose probes with self-alining vanes were unusable on the X-15. Various means of determining these angles were its ar'terbody skin, was selected as the best method, considering the heat transfer, cooling, accuracy, and operational requirements. The feasibility of constructing such a device had been proved, and the components and materials were available. The sensor and its supporting, sealing, and hydraulic-actuating mechanisms are designed as an integral assembly. The electronic amplifiers, power supplies, and control valves are mounted in the afterbody. The electric, hydraulic, and pneumatic connections between the sphere and the cone pass through a single central supporting member. Rotary hydraulic actuators pro- vide the two degrees of freedom required. In operation, the sensor is a null- seeking, hydraulically actuated, electronically controlled servomechanism.

A block diagram of the servomechanization of one axis of the nose sensor is shown in figure 8. Two identical servos are used for independent control in each axis. The differential pressure between opposing orifices is measured, and the unbalance signal is fed through amplifiers to the hydraulic actuator. The actuator then positions the sphere to balance the differential pressures. A synchro transmitter is used to detect the position of the sphere with respect to the airframe, and this signal is f e d to pilot's display instruments, a servorecorder to record the data, and to the telemetry link for ground moni- toring. Since the dynamic pressure c a n vary between 1 lb/sq ft and 2,500 lb/sq ft, compensation is required in the servo loop to maintain stability and accuracy. This compensation is provided by measuring the pressure differ- ence between the total-pressure port a n d one angle-sensing p o r t . The resulting signal is used to adjust the gain of t h e sphere-positioning loop.

T h e angle-of-attack range of the sensor is from -10" to )COO; the angle-of- sideslip range is 9 0 ' .

The unit is capable of continuous operation at a skin temperature of 1,200° F.

Response is flat to about 6 cycles per second with a maximum velocity limit of about 85 deg-/sec. The sensor weighs 78 pounds and is 1 6 3/1 inches long.

The sphere diameter is 6 112 inches, and the base diameter is 13 3/4 inches.

Figure 9 is a comparison of theoretical and actual angle-of-attackmeasure- ments at low dynamic. prcssures. The theoretical angle-of-attackerror (solid -8- -9- curve) was obtained by comparing angle of a t t a c k computed from wind-tunnel and a n a l y t i c a l data a v a i l a b l e f o r t h e sensor and angle of a t t a c k computed from f l i g h t measurements of p i t c h a t t i t u d e obtained from t h e integrated f l i g h t d a t a system and f l i g h t - p a t h angle obtained with t h e precision tracking radar. The a c t u a l angle-of-attack e r r o r was obtained by comparing f l i g h t measurements of angle of a t t a c k from t h e sensor and angle of a t t a c k computed from t h e same i n e r t i a l and radar data. Actual measurements of angle-of-attack e r r o r are shown by t h e s o l i d symbols. The shaded a r e a represents uncertainty i n t h e computation of angle of a t t a c k from i n e r t i a l and radar d a t a . The m a x i m difference i s 3/4" a t a dynamic pressure 3.5 l b / s q f t .

Although it was possible t o measure t o t a l pressure with the nose sensor, t h e l a c k of a s u i t a b l e location on the X-15 f o r a s t a t i c - p r e s s u r e p o r t plus d i f f i c u l t i e s i n sensing and recording t h e low s t a t i c pressures l e d t o t h e s e l e c t i o n of a gyro-stabilized i n e r t i a l reference t o provide a source of A s a result, and because of t h e need f o r measuring v e l o c i t y and a l t i t u d e data.

t r u e a t t i t u d e angles, an integrated i n e r t i a l f l i g h t data system (IFDS) w a s developed. shown i n f i g u r e 10.

A block diagram of t h e system i s The IFDS i s b a s i c a l l y an earth-slaved, schuler-tuned system alined i n azimuth t o an equiva- l e n t guidance equator which i s coincident w i t h t h e radar-range c e n t e r l i n e of t h e X-15. The s t a b i l i z e r u t i l i z e s three self-balancing accelerometers and t h r e e single-degree-of-freedom gyroscopes. A four-gimbal system provides complete a t t i t u d e freedom i n a l l axes. A direct-current analog computer i s used f o r com- puting velocity and position data and the necessary acceleration corrections.

Attitude angles a r e picked d i r e c t l y o f f the gimbals by means of synchro t r a n s - mitters. A Doppler radar i s used a s a v e l o c i t y reference f o r alinement t o t h e v e r t i c a l during prelaunch f l i g h t . A gyro compass i s used f a r a heading r e f e r - ence. The a l t i t u d e loop i s stabilized by s e t t i n g i n the ground g-condition before takeoff and r e f i n i n g t h e s e t t i n g t o c o r r e c t f o r v a r i a t i o n w i t h a l t i t u d e -9- - I U - u n t i l j u s t before launch, when the f i n a l s e t t i n g i s made. A c o n t r o l panel i n t h e B-52 launch a i r c r a f t a i d s i n monitoring t h e alinement process and i n making minor adjustments during c a r r i e d f l i g h t .

The i n e r t i a l system provides t h e measurements shown i n t h e t a b l e below.

The accuracies s e l e c t e d represent a compromise between t h e d a t a desired, t h e X - 1 5 weight and s i z e r e s t r i c t i o n s , and the i n e r t i a l s t a t e of t h e art i n t h e The e i g h t desired measurements may be c l a s s i f i e d i n t o t h r e e 1936-37 period.

groups: (1) a t t i t u d e angles, ( 2 ) a l t i t u d e , and (3) v e l o c i t i e s . The f o u r v e l o c i t i e s are t h e s c a l a r t o t a l or " t r a j e c t o r y " v e l o c i t y and t h e t h r e e compo- nent v e l o c i t y vectors: t h e downrange velocity, t h e crossrange velocity, and t h e The downrange and crossrange v e l o c i t y vectors coincide with v e r t i c a l v e l o c i t y .

t h e v e l o c i t y d a t a obtainable from t h e ground radars.

Measurement s p e c i f i c a t i o n s (Time duration: 300 sec) Measurements required A t t i t u d e angles, deg Unlimit ed 0.5 3,000

Altitude, ft o to 500,000*

Velocity, f t / s e c

- +7, ooo*

Total Downrange +7,000 Cros srange +3,000

Vert i c a1 q, ooo* 20

*Required f o r p i l o t displays.

Figure 1 1 i s a photograph of t h e d i r e c t - c u r r e n t analog computer and t h e Both u n i t s a r e mounted i n t h e main payload com- s t a b i l i z e r , with covers o f f .

partment. The s t a b i l i z e r i s mounted on a s p e c i a l l y constructed v i b r a t i o n i s o l a t o r which minimizes a t t i t u d e changes of t h e instrument with respect t o t h e -10- -11- a i r c r a f t . The computer i s shock-mounted and i s shaped t o conform t o t h e con- t o u r s of t h e payload compartment.

Both u n i t s a r e cooled with cold nitrogen gas t o counteract t h e heat generated by t h e e l e c t r i c a l equipment within t h e u n i t s and heat inputs t o t h e compartment during high Mach number f l i g h t s .

The X-17 inertial-flight-data-system outputs a r e compared with radar data i n f i g u r e 12. Good c o r r e l a t i o n i s evident f o r t h e t o t a l - v e l o c i t y data, and t h e a l t i t u d e curve i l l u s t r a t e s t h e altitude-loop divergence with time t h a t i s t y p i c a l of inertial-guidance systems. This divergence has posed no problem, inasmuch as i n e r t i a l a l t i t u d e data a r e not required f o r f i n a l space position and t h e landing approach.

Aerodynamic-heating measurements.- A s indicated previously, one of t h e primary goals of t h e X-15 program was t o f u r t h e r knowledge of aerodynamic- heating phenomena. This study on the X-15 i s divided i n t o two general research measurement categories: (1) study of s t r u c t u r a l temperatures, including skin temperatures, and ( 2 ) measurement of h e a t - t r a n s f e r r a t e s t o t h e a i r p l a n e .

During t h e instrumentation design phase: it soon became apparent t h a t t h e only f e a s i b l e method a v a i l a b l e f o r making these measurements involved t h e use of thermocouples attached t o t h e skin and i n t e r n a l s t r u c t u r e plus t h e use of surface pressure measurements t o determine l o c a l flow conditions. The use of thermocouples on t h e X-15 required considerable development and t e s t work t o obtain s u i t a b l e thermocouple materials, t o devise a recording scheme that would a g r e a t number of thermocouple outputs t o be recorded on each f l i g h t , and, allow f i n a l l y , t o decide upon a thermocouple-attachment method that would provide p r o t e c t i o n f o r t h e thermocouple during periods of high temperature and high vtbration. I n addition, t h e thermocouples would be inaccessible a f t e r t h e a i r p l a n e was constructed; hence, the i n s t a l l a t i o n would need t o be long-lived with l i t t l e o r no maintenance.

-11- -12 - Figure 1 3 i l l u s t r a t e s t h e i n s t a l l a t i o n of a t y p i c a l s t r u c t u r a l thermo- couple. The thermocouple m a t e r i a l selected was 30-gage chromel-alumel. The 30-gage l e a d s a r e spot-welded t o t h e structure and routed t o 20-gage thermo- couple extension l e a d s . The use of 20-gage extension l e a d s t o t h e signal con- ditioning, reference junction, and recording equipment was necessary t o reduce c i r c u i t r e s i s t a n c e i n t h e thermocouple loops and t o minimize measurement e r r o r s due t o r e s i s t a n c e changes caused by l a r g e temperature v a r i a t i o n s along t h e wire. The thermocouple l e a d s are insulated from each other by s i l i c o n e - impregnated Fiberglas b r a i d . Outer i n s u l a t i o n i s t h e same. Where t h e thermo- couple l e a d s are c l o s e t o hot s k i n or o t h e r s t r u c t u r e , a n o u t e r sleeve of unimpregnated Fiberglas sleeving g i v e s a d d i t i o n a l protection. The length of thermocouple l e a d enclosed i n t h e sleeving i s held firmly t o t h e member by w i r e tiedowns spot-welded t o t h e member t o minimize v i b r a t i o n and shock damage t o t h e i n s u l a t i o n . The s i l i c o n e impregnation gives s t r e n g t h t o t h e Fiberglas i n s u l a - t i o n , thus allowing it t o withstand t h e s t r e s s e s of i n s t a l l a t i o n . The impreg- nation eventually sublimates during repeated exposure t o elevated temperatures b u t maintains i t s e l e c t r i c a l i n s u l a t i n g p r o p e r t i e s . Tests indicated t h a t t h i s gassing-off process could r e s u l t i n a n explosion i f t h e sample were suddenly brought t o 1,200° F. The hazard w a s eliminated on t h e X-13 by a gradual buildup t o maximum Mach number (1,200° F) during t h e course of t h e program.

A n i n t e r n a l v i e w of a t y p i c a l X-15 skin thermocouple and surface-pressure- p o r t i n s t a l l a t i o n i s shown i n f i g u r e 14. The surface-pressure o r i f i c e s are f i x e d i n place by a simple clamp designed t o prevent any r e l a t i v e movement between t h e o r i f i c e and wing skin. Tubing of l/b-inch i n n e r diameter, 0.025-inch w a l l thickness,fabricated of seamless Inconel and s t a i n l e s s s t e e l i s used t o route t h e sensed pressure t o t h e NASA precision pressure recorders i n t h e instrument compartments. These tubing m a t e r i a l s were u t i l i z e d t o provide The inner diameter and w a l l thickness were s t r e n g t h a t high temperatures.

-12- -13- chosen as a compromise which would r e s u l t i n minimum pressure l a g and minimum space and weight.

Figure 13 i s a view of t h e thermocouple s i g n a l conditioner. The condi- t i o n e r c o n s i s t s of a mechanical commutator capable of switching a t o t a l of 480 thermocouples onto 12 oscillograph channels a t a r a t e of 40 per channel p e r second, t h e thermocouple reference junction, and t h e necessary i n - f l i g h t zero and c a i i b r a t i o n provisions. The thermocouple reference junctions a r e maintained a t constant temperature.

Data from a t y p i c a l X - 1 5 h e a t - t r a n s f e r f l i g h t a r e shown i n f i g u r e 16.

Three thermocouple outputs a r e compared with calculated temperatures; good c o r r e l a t i o n i s evident f o r a l l three l o c a t i o n s . The divergence of measured and c a l c u l a t e d d a t a f o r t h e spar-web thermocouple i l l u s t r a t e s t h e d i f f i c u l t y i n c o r r e l a t i n g t h e o r e t i c a l calculations and measured s t r u c t u r a l temperature d a t a .

GROUND INSTRUMENTATION Ground Range One of t h e b a s i c instrumentation requirements f o r a highly experimental program such as t h e X - 1 5 i s a well-instrumented ground range. The ground range f o r t h e X - 1 5 i s shown i n f i g u r e 17 i n r e l a t i o n t o a t y p i c a l X - 1 5 mission. The range i s required t o perform t h e following s p e c i f i c functions during an X - 1 5 f l i g h t : Aid i n t h e i n i t i a l guidance and vectoring of t h e B-52 launch a i r p l a n e t o t h e required heading.

Monitor t h e i n i t i a l climb of t h e X - 1 5 a i r p l a n e .

Provide a backup f o r a l t i t u d e and v e l o c i t y information t o t h e p i l o t i n t h e event of on-board equipment f a i l u r e .

Monitor t h e f l i g h t path as an aid i n homing or vectoring t o a s u i t a b l e intermediate emergency landing area, i f required.

-13- -14-

Provide -

Information f o r e s c o r t airplane rendezvous.

F i n a l approach and landing information t o t h e p i l o t .

Reliable long -range c o m m i c a t i o n s c a p a b i l i t y .

A real-time data-monitoring c a p a b i l i t y .

Accurate space-trajectory d a t a f o r research purposes.

To meet these requirements, a ground range w a s constructed c o n s i s t i n g of t h r e e s t a t i o n s : a main terminal s t a t i o n i n t h e N A S A F l i g h t Research Center a t Edwards, C a l i f . , and two up-range s t a t i o n s , one i n t h e v i c i n i t y of Beatty, Nev., t h e other near Ely, Nev. Many considerations entered i n t o t h e choice of t h e s p e c i f i c l o c a t i o n s f o r t h e up-range s i t e s , including required r a d a r overlap c a p a b i l i t i e s , t h e power balance i n t h e radar-to-beacon loop, t h e requirement of a m a x i m u m omnidirectional seeing angle, and t h e o v e r a l l l o g i s t i c problem. The l o c a t i o n s and e l e v a t i o n s of t h e s i t e s are such that omnidirectional tracking can be accomplished down t o an a l t i t u d e o f 10,000 f e e t .

A f u n c t i o n a l diagram of t h e X-15 range i s shown i n f i g u r e 18. The radar system includes t h e r a d a r s and a l l a u x i l i a r y equipment necessary t o provide s e q u e n t i a l tracking of t h e t e s t a i r c r a f t by t h e t h r e e s t a t i o n s from one end of t h e range t o t h e other. The radar i s an automatic angle and range tracking u n i t designed t o provide accurate azimuth and e l e v a t i o n angle and s l a n t range d a t a . These u n i t s operate i n "S" band and have a 400-mile ranging c i r c u i t c a p a b i l i t y . The radar may be positioned by means of a remote o p t i c a l t r a c k e r or by computed analog d a t a from t h e radar data-acquisition system. These remote i n p u t s a r e used for t a r g e t a c q u i s i t i o n and as t r a c k i n g a i d s .

A p r e c i s i o n data-recording system operates i n conjunction with t h e r a d a r t o yermit continuous Trisual monitoring of t.arget space p o s i t i o n (azimuth, elevation, and range) and r e a l time of day, and provides photographic and A n 80-inch magnetic-tape recordings of t h e t a r g e t position and time of day.

-14-

-15 -

focal length boresight camera associated with the system permits photography of the airborne target at limited ranges. Tracking information in the f o r m of azimuth, elevation, and range is obtained from two optical digital encoders and one electromechanical encoder (range) which are attached directly to the radar antenna and range shafts. Digital information from the encoders is recorded on magnetic tape and constitutes the primary information obtained from the radar system. In addition to the tape recorder, a data camera is The camera photographs the same digital infor- included for backup purposes.

mation as recorded on the magnetic tape and selsyn dial indications of azimuth, elevation, and range for coarse trajectory information.

Radar data are transmitted between stations f o r plotting purposes as well as for radar pre-position inputs to aid in sequential tracking of the X-15 Analog voltages representing space position and velocity over the entire range.

are converted to digital form and transmitted by either telephone lines or microwave to the next station on the range.

The telemetry ground-receiving, conditioning, and recording equipments are a standard 18-channel FM-FM system and a pulse-duration-modulation system capable of receiving up to 90 channels of information. Reliable reception is The achieved through use of a servo-driven cross-dipole tracking antenna.

antenna can be controlled by synchro data from the tracking radar so that the Complete data from both the telemetry antenna automatically tracks the target.

FM-FM and pulse-duration-modulationstations are recorded on magnetic tape.

Telemetry data from the up-range stations, Beatty and Ely, are transmitted via microwave to the main terminal station at Edwards, recorded on magnetic tape, and displayed in real time.

Voice communications with the X-15 is accomplished by means of standard Communication with support aircraft and ground military UHF ground equipment.

vehicles is accomplished through the use of UHF equipment and single-sideband high-frequency equipment. When a UHF transmitter, or transmitters, is keyed at -15- -16 - any range s t a t i o n , t r a n s m i t t e r s a t t h e other two s t a t i o n s a r e keyed by means of s i g n a l s sent over t h e microwave o r telephone l i n e s , s o that t h e same information i s transmitted simultaneously by a11 three s t a t i o n s . Conversely, UHF receivers a t a l l t h r e e s t a t i o n s feed t h e i r outputs i n t o t h e microwave o r hardlines so that t h e information i s received and heard a t a l l s t a t i o n s . Undesirable b e a t f r e - quencies, which may be caused by transmitter c a r r i e r d r i f t , a r e eliminated by o f f s e t t i n g t h e c a r r i e r frequencies of the t r a n s m i t t e r s a t each of t h e t h r e e s t a t i o n s . A s t a t i o n - t o - s t a t i o n intercommunications system u t i l i z i n g t h e micro- wave and telephone l i n e s i s available for range administrative and control purposes.

A timing system furnishes precise timing reference s i g n a l s t o t h e t h r e e s t a t i o n s f o r t h e purpose of c o r r e l a t i n g data recorded by t h e various instrumen- t a t i o n f a c i l i t i e s a t t h e individual s i t e s . The timing reference s i g n a l s a r e transmitted by telephone l i n e s .

Figure 19 i s a photograph of t h e X-17 ground c o n t r o l room i n t h e main terminal s t a t i o n a t Edwards. I n t h i s room a r e t h e radar p l o t t i n g boards and t h e monitor consoles a t which engineers responsible f o r each subsystem i n t h e X-15 may monitor c r i t i c a l parameters i n real time. These displays t a k e t h e form of oscilloscope b a r c h a r t s f o r l i m i t displays, meter presentations, and s t r i p - c h a r t time h i s t o r i e s .

Flight Simulator A major r o l e i n t h e X-13 ground instrumentation complex has been played The simulator c o n s i s t s of an by t h e X-15 analog f l i g h t simulator ( f i g . 2 0 ) .

e l e c t r o n i c analog computer t o solve t h e equations of motion i n s i x degrees of freedom and a fixed-base cockpit and control-system mockup. The simulator i s used f o r d e t a i l e d t r a j e c t o r y planning, p i l o t t r a i n i n g f o r normal and emergency conditions, exploration of c r i t i c a l s t a b i l i t y and c o n t r o l areas, extrapolation -16 -

-17 -

of a c t u a l X-15 c h a r a c t e r i s t i c s t o unexplored areas, and i n v e s t i g a t i o n i n t o e f f e c t s of proposed modifications t o t h e X-17 systems or configuration.

A block diagram of t h e analog simulator i s shown i n f i g u r e 21. The cockpit as c l o s e l y as possible, every d e t a i l of t h e X - 1 5 cockpit, simulator duplicates, The hydraulic c o n t r o l system has been including instrumentation and controls.

flow rates, mechanized t o duplicate t h e a c t u a l hydraulic l i n k s and pressure and as well as t h e mass c h a r a c t e r i s t i c s and n a t u r a l freqEencies of t h e X - 1 7 h o r i - Control-system nonlinearities a r e a l s o c l o s e l y duplicated.

z o n t a l t a i l .

Two s t a b i l i t y augmentation systems a r e used on t h e X - 1 5 airplane, as shown One i s a fixed but selectable gain system (SAS) and t h e other, i n f i g u r e 21.

These systems a r e included on t h e simulator through adaptive or variable gain.

t h e use of a c t u a l vehicle systems electronic u n i t s and a r e made compatible with t h e computer with signal-conditioning elements. The simulator a l s o includes a malfunction generator which simulates the major systems f a i l u r e s t h a t might occur i n f l i g h t . Simulated f a i l u r e s of any of the major X - 1 5 cockpit i n s t r u - ments or 23 possible systems f a i l u r e s a r e indicated i n t h e simulator cockpit by l i g h t s .

The data obtained from the airborne recorders a r e processed a t t h e F l i g h t Research Center with an IBM 704 computer, The raw data on t h e oscillograph and photorecorder f i l m s a r e transferred t o IBM punched cards by using manually operated f i l m readers. Computer input tapes a r e then prepared. Magnetic tapes from t h e range s t a t i o n s a r e processed automatically by the d i g i t a l data- processing u n i t of t h e U , S . A i r Force a t Edwards. Radar tapes a r e converted t o IBM computer input tapes from which geometric a l t i t u d e , plan position, t r a - jecltory position, and velocity a r e obtained and used f o r research analysis, comparison with data derived f r o m t h e i n e r t i a l data system, and operational a n a l y s i s of radar performance.

Telemetry tapes may be played back on t h e -17 - -18- Edwards ground s t a t i o n f o r quick-look purposes o r processed by t h e A i r Force u n i t f o r more p r e c i s e r e s u l t s .

OPERATIONAL EXPERIENCE Since t h e start of t h e X - 1 5 f l i g h t program i n t h e spring of 1939, t h e a i r p l a n e has flown t o s l i g h t l y over Mach 6 and exceeded an a l t i t u d e of 350,000 f e e t . The r i g o r s imposed upon the airborne instrumentation components and systems by sustained operations i n t h e envelope bounded by t h e s e limits have demonstrated t h a t t h e philosophy adopted t o meet t h e program i n s t m e n - t a t i o n requirements w a s adequate. The r e l i a b i l i t y of t h e t h r e e major X - 1 3 instrumentation systems i s indicated i n t h e following t a b l e : R e l i a b i l i t y , Data System Flights percent 100 Acceptable Airborne data a c q u i s i t i o n 9 7 N A S A hypersonic nose a, B, P ' 70 98 acceptable sensor Integrated i n e r t i a l A c c e ptab 1 e f l i g h t data A s shown i n t h e t a b l e , t h e airborne data-acquisition system has f u l f i l l e d The r e l i a b i l i t y of t h e system w a s a l l t h e requirements imposed upon it.

c a l c u l a t e d by dividing t h e t o t a l number of usable d a t a channels by t h e t o t a l number of required d a t a channels f o r 100 f l i g h t s of t h e t h r e e X-15 a i r c r a f t .

This percentage includes t h e data-recording equipment and d i s p l a y f o r t h e i n e r t i a l f l i g h t data system and t h e nose sensor.

No major performance problems have been encountered with t h e airborne The senscr i n s t a i i a t i o n s i n hot a r e a s of t h e a i r c r a f t d a t a - a c q u i s i t i o n system.

have met design objectives and have withstood t h e high temperatures and -18- -19- v i b r a t i o n l e v e l s encountered. It has not been possible t o s t a y within t h e o r i g i n a l weight c o n s t r a i n t imposed on t h e system; t h e t o t a l weight of t h e sys- tem including t h e nose sensor and i n e r t i a l system i s 1,400 pounds. The data- processing problem has developed as anticipated. Although t h e method used i s not s a t i s f a c t o r y i n t h e l i g h t of today's s t a t e of t h e art, it has f u l f i l l e d pro- gram requirements. The number of data points processed per f l i g h t has averaged between 8,000 t o 10,000, and tabulated data have been given t o t h e f l i g h t - t e s t The most d i f f i c u l t t a s k engineers approximately one week a f t e r f l i g h t .

encountered i n processing t h e data from t h e airborne system was i n t r a n s f e r r i n g t h e thermocouple data from t h e oscillograph f i l m t o punched IBM cards. A t first t h i s was a time-consuming, tedious job but as operator proficiency increased t h e time required was reduced t o acceptable l e v e l s .

The NASA hypersonic nose sensor has had a remarkable r e l i a b i l i t y record.

sensor has been used, it operated marginally on O f t h e 7 0 f l i g h t s on which the only 1 f l i g h t .

The integrated i n e r t i a l f l i g h t data system has been used i n 89 X-13 The r e l i a b i l i t y shown i n the preceding table indicates t h e approxi- f l i g h t s .

T h i s percentage i s based on t h e number mate o v e r a l l r e l i a b i l i t y of t h e system.

an i n - f l i g h t of times t h e system has caused a major delay i n a scheduled f l i g h t , abort p r i o r t o launch, or has failed during f l i g h t . The system, however, has not been a b l e t o perform consistently within the performance s p e c i f i c a t i o n s s e t i n 1956, I n retrospect, t h e performance s p e c i f i c a t i o n s established a t t h a t t i m e were beyond t h e c a p a b i l i t i e s of t h e s t a t e of t h e a r t w i t h respect t o avail- able gyros, accelerometers, t r a n s i s t o r s , and c i r c u i t techniques. However, t h e system has been a b l e t o perform a t l e v e l s which, although marginal o r subpar i n regard t o t h e specifications, have enabled t h e f u l l performance c a p a b i l i t i e s of t h e X-15 t o be r e a l i z e d . Continued development, through hardware m o d i f i - cations as new c i r c u i t techniques, e l e c t r o n i c components, and improved gyros -19- -20- became available, plus refinements to operational and quality-control proce- dures have steadily improved the system as the flight program has progressed.

A new system, based upon technology and hardware developed by the U.S. Air Force for the X - 2 0 program, is being procured to replace the present system.

FOLLOW-ON PROGRAM The X - l 5 flight program is at the stage where much of the flight research required to satisfy the original goals has been achieved. The three X - 1 5 airplanes are now embarked in a follow-on program for flight testing scientific instruments, propulsion systems, and advanced flight data and energy management systems, and for studies of the atmosphere and advanced structures research.

Included in these tests are: ultraviolet stellar photography, ultraviolet and infrared exhaust-plume observation, horizon definition, optical-degradation measurements, high-temperature-window evaluations, high-temperature leading-edge studies, atmospheric-densitymeasurements, micrometeorite collection, and evalu- ations of an advanced integrated flight data system, energy management, vapor- cycle cooling, and airbreathing propulsion systems. The data-acquisition system used during the basic program w i l l support these tests.

CONCLUDING REMARKS Many valuable instrumentation lessons applicable to future vehicles have been learned from the X-13 program. Among the most important are the following: The oscillographic data-acquisition system used in the X-15 repre- sents about the ultimate in the application of this type of system in regard to number of channels, weight per channel, accuracy, and the data-processing task. Although the system has done its job, if the same ground rules were used today as in 1957, a digital, magnetic- tape system with automatic data processing would be selected.

-20- -21- A system such as an FM-FM magnetic-tape recording system f o r t h e measurement of dynamic phenomena such as f l u t t e r , vibration, and acoustic noise should always be included i n t h e o r i g i n a l instrumen- t a t i o n layout on a vehicle such as t h e X-13. This recorder was found t o be a necessary addition t o t h e X - 1 5 data-acquisition system during t h e f l i g h t program.

The necessity f o r f l i g h t t e s t i n g of complex, new instrument systems before they a r e required t o produce data on a f l i g h t program similar t o t h e X - 1 5 i s imperative i n order t o b u i l d confidence and discover deficiencies e a r l y enough f o r c o r r e c t i v e measures t o be applied. This was amply demonstrated by t h e X-13 i n e r t i a l system, which w a s inadequately f l i g h t t e s t e d before i t s use was required.

Low-level, low-speed electronic commutators have proved t o be r e l i a b l e on X-15 f l i g h t s , obviating the need f o r using low-level, mechanical commutators with t h e i r attendant contact noise and main- tenance problems.

-22 -

SYMBOLS a l t i t u d e , f t Mach number surface pressure nose sphere, l b / s q f t d i f f e r e n t i a l pressure, lb/sq f t t o t a l pressure, l b / s q f t s t a t i c pressure, lb/sq f t dynamic pressure, lb/sq f t time, sec acceleration along X, Y, and Z body axes, respectively, g u n i t s t o t a l s c a l a r v e l o c i t y ( i n e r t i a l ) , f t / s e c downrange v e l o c i t y ( i n e r t i a l ) , f t / s e c v e r t i c a l v e l o c i t y ( i n e r t i a l ) , f t / s e c crossrange v e l o c i t y ( i n e r t i a l ) , f t / s e c angle of attack, deg angle-of-attack e r r o r , deg angle of s i d e s l i p , deg f l i g h t - p a t h argle, deg angle of pitch, deg angle of roll, deg angle of yaw, deg

-22 -

X-15 PaKORMANCE €NEW 350 IO3

\

DEWN ALTlTUDE e__--

250 -

200 -

ALTITUDE, FT

150 -

100 -

50 -

VELOCITY, FT/SU: Figure 2

TYPICAL X-15 MISSIONS

Figure 3

X-15 AIRPLANE SURFACE INSTRUMENTATION

293 THERMOCOUPLES 136 PRESSURE ORIFICES * THERMOCOUPLES + PRESSURE ORIFICES Figure 4

X-I 5 INSTRUMENT COMPARTMENTS

CENTER OF GRAVITY MAIN INSTRUMENTATNM NOSE X-15 INSTRUMENTATION SYSTEM NOSE wy.

TAIL CE"ER-OF-GRWITY COMPARTMENT HIPERSQlK mYPARTYENT COMPARTMENT NOSE SENSOR EVENTS (1. B FORCES PHYSIOLOGICAL a. B -PILOT'S PRECISKIN NASA DISFIAY RECORDER h ~ ~ ~ e * * * v * v v h.V. V,.V,.V, I I TELEMETRY INTEGRATED IN E R T I AL FLIGHT RECORDER R E C W D E R ~ I SYSTEM

NASA SENSOR INSTALLED ON THE X-15

AFT CONE REMOVED SENSOR MECHANIZATION a OR / 3 AXIS HYDRAULIC PCTUmOR TO INDICATOR A N D R E C M D E R Figure 8 COMPARISON OF ANG LE-OF-ATTACK MEASUREMENTS AT LOW DYNAMIC PRESSURE ATTACK ERROR, DEG 2 3 4 5 I O 20 30 40 WNAMtC PRESSURE, LB/so FT INTEGRATED INERTIAL FLIGHT D A T A SYSTEM MAJOR 8-52 EQUIPMENT

I

I

MODE CONTROL DOPPLER RADAR SYSTEM MONITOR REFERENCE DATA COMPASS SYSTEM *-- ? UMBILICAL CONNECTOR ( .-_' MAJOR X-15 EQUIPMENT

I PILOT'S DISPLAY I

*

I /CObP*] ATTITUDES

I I

TOTAL VELOCITY VERTICAL VELOCITY ALTITUDE Figure 10

IFDS COMPUTER AND STABILIZER

CHARACTERISTICS OF INERTIAL FLIGHT DATA SYSTEM VELOCITY, FT/SEC I 2 0 ALTITUDE, FT 0 160 320 480 640 TIME, SEC Figure 12 X-15 SKIN THERMOCOUPLE INSTALLATION SILICONE-IMPREGNAlED

d FIBERGLAS INSULATION

FIBERGLAS SLEEVING WING SKIN UNIMPREGNATED IN ' CONTACT WiTH SKIN HOT JUNCTION WIRE TIEDOWNS

e - TO SIGNAL

COlSMTlONR Figure 13

TYPICAL X-15 THERMOCOUPLE AND

PRESSURE-fWtT INSTALLATION

.

X-15 THERMOCOUPLE SIGNAL COMDITIONER

COMPARISON OF CALCULbTED AND MEASURED TEMPERATURES HEAT-TRANSFER FLIGHT - I 200 1000 - 800 - - TEMPERATURE, O F 400 -

200 c

TIME, SEC Figure 1 6

TYPICAL X-15 RESEARCH MISSION

BURNOUT - 1 1.85 SEC h=158DOO FT / V = 5,460 FT/S€C

I

, RANGE FUNCTIONAL DIAGRAM RADAR, TELEMETER, VOICE .

LOCAL MICROWAVE AND TELEPHONE LINE INTERCONNECTING CIRCUITS m RECORDED AT EACH STATIO) TRANSMITTFD BFTWFFN S TATIONS PRECISION RADAR DATA RADAR ACQUISITION DATA TEEMETERING VOCE VOCE TIMING TELEW3RV W A Figure 18

X-15 GROUND CONTROL STATION AT FRC

.

Figure 19 FIXED-BASE SIMULATOR AIRPLANE MOCKUP COMRETE COU(FIT &7ML HARMARE CONTROL SYSTEM KTUAL ELECTRONIC SAS 3 MOOEL 231R EA ANALOG COMPUTERS 92 W E FUNCTION GENERATORS 33 COMPUTING SERVOS 5 ELECTRONIC MULTIPLIERS Figure 20 ANALOG-SIMULATOR DIAGRAM STABILITY AND CONTROL AUGMENTATION SYSTEMS I r------i I A I 1 I MALFUNCTION GENERATCR CONTROL SYSTEM COCKPIT SIMULATOR

I- 1

I I ------ ------ I

Figure 21