summary of high-altitude and entry flight control experience with the

NASA TECHNICAL NOTE

SUMMARY OF HIGH-ALTITUDE AND ENTRY FLIGHT CONTROL EXPERIENCE WITH THE X-15 AIRPLANE

by Ed,id C. Holleman r

Flight Research Center Eawaras, cdli~;

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N A T I O N A L AERONAUTICS AND SPACE A D M I N I S T R A T I O N WASHINGTON, 0. C. . APRI? 1966

https://ntrs.nasa.gov/search.jsp?R=19660011752 2018-02-15T22:15:32+00:00Z

TECH LIBRARY KAFB. NM

SUMMARY O F HIGH-ALTITUDE AND

ENTRY FLIGHT CONTROL EXPERIENCE WITH

THE X-15 AIRPLANE

By Euclid C. Holleman

Flight Research Center Edwards, Calif.

N A T I O N A L AERONAUTICS AND SPACE ADMINISTRATION

0330234

For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $1.00

I

SUMMARY OF HIGH-ALTITUDE AND

EWlBY FLIGHT CONTROL EXPERIENCE: WITH

By Euclid C. Holleman Flight Research Center

Flights to high altitude with the X-15 research airplane have successfully demonstrated that a pilot can control lifting entry into the atmosphere. Flights were made with two airplane configurations and with several types of reaction and aerodynamic controls.

The pilots' performance in these flights was similar with all of the reaction control systems tested--angular acceleration, angular acceleration with damping, and rate command--and was comparable to the performance of the automatic hold systems. The pilots evaluated the reaction control task with all systems as satisfactory. For a range of altitude from 200,000 feet to 354,200 feet, the entry control task with attitude command with damping, rate command, and hold was rated satisfactory. The entry acceleration environment did not affect the pilot's performance. Neither the controllability of the entry nor the pilot's evaluation of the task was degraded by the more severe entries.

With the preflight procedures established for the X - 1 5 program, control of range to the landing site has been accomplished easily by the pilots. Flights to high altitude have been planned with 50 to 100 nautical miles of excess range, and the pilots have been able to judge and control excess range by modulating angle of attack and using speed brakes to insure landing at the desired landing site.

INTRODUCTION

The X-15 airplane is the first piloted airplane designed for atmospheric entry and hypersonic flight research. As such, it is capable of investigating flight problems at extremely high altitudes and at high speeds in an actual entry flight environment characterized by high acceleration forces and both rapid and large changes in airplane response characteristics.

In a flight program designed to explore the entry control characteristics of the X-13 airplane, 12 flights to approximately 200,000 feet or greater were

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made, thus providing piloted entry experience from high altitude. plane configurations and several different reaction and stability augmentation controls were used.

Two air-

This paper summarizes the high-altitude X-15 flight experience, which culminated in a flight to an altitude of 354,200 feet. stability, control, and handling characteristics of the airplane, the cockpit displays, and the operational techniques that enabled it to be successfully flown to and recovered from high altitudes without special piloting aids other than stability augmentation. with the airplane equipped with interim rocket engines is discussed in reference 1.

Discussed are the basic

Flight experience to moderately high altitudes

The symbols used in this paper are defined in the appendix.

AIRPLANE AND SYSTEMS

Airplane

The X-15 is a single-place, rocket-powered airplane (figs. 1 and 2) designed for flight at hypersonic speeds and extreme altitudes. It is carried aloft under the right wing of a B-52 aircraft and is launched at an altitude of about 45,000 feet and a Mach number of about 0.80. performs a powered flight mission followed by a deceleration glide before vectoring for a landing at Edwards Air Force Base, Calif. tional technique, the airplane is capable of a Mach number of 6 and can be flown to and recovered from an altitude in excess of 3OO,OOO feet.

After launch, the X-15

With this opera-

Flights to high altitude have been made with two airplane configurations-- lower movable ventral on, and lower ventral off (fig. 2). control and physical characteristics of these configurations are presented in detail in reference 2. The static-stability derivatives of the basic airplane (lower ventral on) indicate that the vehicle should be stable throughout the flight envelope; however, near a Mach number of 3.0 and an angle of attack of 10" the airplane was uncontrollable without damping augmentation and with the pilot controlling bank angle normally. An analysis of this instability re- vealed the cause to be an unfavorable combination of the yawing moment due to aileron deflection and the dihedral effect, which was subsequently alleviated by removing the lower movable ventral (ref. 3). The change, although producing lower static-directional stability and rudder effectiveness, resulted in a more controllable airplane with damping augmentation inoperative, particularly at high angles of attack.

The stability and

Aerodynamic control is provided through conventional aerodynamic surfaces; however, the horizontal tail provides both pitch and roll control. The aerodynamic control surfaces are actuated by irreversible hydraulic systems. Artificial feel is provided for control feel. A conventional center stick, a right-side-located controller, and rudder pedals provide aerodynamic control. The controller, although designed for use in high-acceleration environments, was used by the pilots throughout the aerodynamic flight regime. A

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left-side-located three-axis controller is also provided for use with the reaction control system. This control system is described in detail in reference IC.

Augmentation Systems

To provide adequate handling qualities over the aerodynamic operating envelope of the X-15 airplane, aerodynamic damping augmentation about all three axes is necessary. The two systems designed for this purpose are a conventional stability augmentation (damper) system, referred to as SAS, and the adaptive flight control system, referred to as AFCS. Detailed descriptions of these systems may be found in references 4 and 5 (SAS) and 6 to 8 (AFCS) .

Stability augmentation system.- The SAS provides auxiliary aerodynamic damping by actuating the aerodynamic control surfaces to oppose the rotational velocity of the airplane. For the basic airplane configuration, an inter- connect damper loop (termed yar) provides a crossfeed yaw-rate signal into the roll control surfaces. With the ventral-off configuration, the yar inter- connect is undesirable and, therefore, not used. Because of a need for redundancy, a backup augmentation system has been designed and installed in the roll modes of the two X-15 airplanes that are equipped with stability augmentation systems.

Adaptive flight control system.- The m C S used in the other aircraft is a model-following, rate command system. The principal features of the system are: self-adjusting gains, rate-command control by the pilot, hold or attitude command modes of operation, normal-acceleration command and limiting, and automatic blending of aerodynamic and jet reaction controls. Most of the high- altitude flights were made with this system because of its inherent high performance, fail-safety features, and high reliability.

Reaction Controls

Four modes of operation--angular acceleration command with and without reaction augmentation damping, angular rate command, and attitude command or hold--are available with the reaction control system (see refs. 6 to 9). the modes have been used during flights to extremely low dynamic pressure. The basic reaction control system is dualized for redundancy. Dam-ping augmentation has been provided by paralleling one channel of the basic system with an electronically controlled rate damper and is available only when reaction control thrust is not commanded by the pilot.

All

The reaction controls for the airplane equipped with the adaptive flight control system may be operated in any combination of three modes: the basic angular-acceleration-command mode activated by the separate three-axis controller, an angular rate command mode integrated with the aerodynamic controls, and an attitude hold or attitude command mode also integrated with the aerodynamic controls.

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Displays

A photograph of the p i l o t ' s display i s shown i n f igu re 3. The three-axis a t t i t u d e ind ica to r (center ) i s the primary instrument f o r displaying airplane p i tch , roll, and yaw a t t i t u d e t o the p i l o t . A p i t ch -a t t i t ude vern ier i s provided on the l e f t s ide of the ind ica tor f o r more accurate d isp lay of p i t c h a t t i t u d e . A d isp lay giving dynamic pressure i s located above the a t t i t u d e indicator , and immediately above the dynamic-pressure d isp lay i s a t imer which displays elapsed engine t h r u s t time. Early i n the program when i n e r t i a l ve loc i ty w a s not y e t considered t o be r e l i ab le , t h r u s t t i m e w a s used as a cue t o shut down the engine.

Angle of a t t a c k and s i d e s l i p a re presented on the cross bars of the three- ax i s a t t i t u d e ind ica tor . Actual readings of a and p a re not given on the ind ica tor , but the angle-of-attack bar may be s e t as a vernier for es tab l i sh ing en t ry angle of a t t ack . A t high a l t i t u d e where the s i d e s l i p ind ica t ion i s unrel iable , the s i d e s l i p bar may be switched t o present heading change. Angle of a t t ack i s a l s o presented on a d i a l gage t o the l e f t of the a t t i t u d e ind i - ca tor . Above the angle-of-attack d i a l i s a presentat ion of normal acceleration.

Conventional pressure-derived airspeed and a l t i t u d e a r e displayed t o the l e f t of the angle-of-at tack d i a l f o r use a t Mach numbers l e s s than 2 and a l t i - tudes l e s s than 80,000 f e e t . Velocity, a l t i t u d e , and r a t e of climb, derived from the i n e r t i a l reference platform, a re displayed i n the upper-right corner of the panel. A display of r o l l r a t e w a s included as a secondary display f o r use i n the event the augmentation system f a i l e d .

The X-15 a i rp lanes a re operated only during contact f l i g h t conditions, so no guidance displays, with the exception of heading, a r e provided for the p i l o t .

OPEFUITIONAL PLAN

The philosophy of the X-15 f l i g h t - t e s t program, as s t a t ed i n reference 10, was t o expand the f l i g h t envelope t o the maximum speed and a l t i t u d e as rapidly as p r a c t i c a l and t o obtain, simultaneously, as much research da ta i n hypersonic f l i g h t as possible . The envelope-expansion program has been performed on an incremental performance bas i s ; t h a t i s , each successive f l i g h t i s designed t o go t o a s l i g h t l y higher speed o r a l t i t u d e than the previous f l i g h t . This approach permits a reasonable extrapolat ion of f l i g h t - t e s t da ta from one f l i g h t t o the next and bui lds a backlog of p i l o t experience.

The f l i g h t plan f o r a t y p i c a l a l t i t u d e mission i s shown i n t ab le I, and the time h i s to ry of a f l i g h t t o ah a l t i t u d e of 285,OOO f e e t i s shown i n f i g - ure 4. t h r u s t and r o t a t e s t he airplane u n t i l the l imi t ing condition of 2g or 10" of angle of a t t ack i s reached. e x i t p i t c h angle (42") i s a t t a ined . Thereafter, t he p i t c h vernier on the three-axis b a l l allows the p i l o t t o f l y a constant p i t c h a t t i t u d e t o within 1". A t the extreme p i t c h angles, the p i l o t must r e l y e n t i r e l y on h i s displays,

Immediately a f t e r launch, the p i l o t opens the t h r o t t l e t o 100-percent

This condition i s maintained u n t i l the desired

s ince contact f l i g h t i s inaccurate and impractical . The iner t ia l -system indica t ions of ve loc i ty and a l t i t u d e and the radar a l t i t u d e ca l lou t s from the ground provide the p i l o t with addi t iona l cues. down time and e x i t p i t c h a t t i t u d e a re the two quan t i t i e s t h a t the p i l o t m u s t cont ro l e f f ec t ive ly during the powered port ion of t he f l i g h t i n order t o con t ro l maximum a l t i t u d e .

F ina l ve loc i ty or engine shut-

After engine shutdown during the low-dynamic-pressure port ion of the f l i g h t , the p i l o t uses the reac t ion controls , on cue from the three-axis a t t i t u d e indicator , t o cont ro l the airplane a t t i t u d e . The airplane i s i n a b a l l i s t i c t r a j e c t o r y and only a t t i t u d e i s cont ro l lab le . The desired en t ry angle of a t t ack i s es tab l i shed by using the reac t ion cont ro ls . As dynamic pressure bui lds up, angle of a t t ack i s maintained u n t i l normal accelerat ion increases t o the l e v e l required f o r pul lout . This acce lera t ion i s held u n t i l l e v e l f l i g h t i s achieved and a g l ide t o the landing s i t e i s es tabl ished. The pressure instruments ( a l t i t u d e , airspeed) a re used a f t e r the airplane reaches Mach numbers l e s s than 2 and a l t i t u d e s l e s s than 80,000 f e e t t o perform the approach and landing.

P i l o t Preparation

Before each f l i g h t , the p i l o t p rac t ices the proposed f l i g h t plan using the six-degree-of-freedom X-15 fixed-base simulator. Both the p i l o t and the ground con t ro l l e r become fami l i a r with the required p i l o t i n g techniques and timing of t he proposed f l i g h t plan. The desired f l i g h t i s "flown" severa l times, and changes suggested by the p i l o t a r e incorporated i n t o the f l i g h t plan. After t he p i l o t i s thoroughly familiar with the f l i g h t plan, f l i g h t s s l i g h t l y off the desired a re flown t o acquaint him with the e f f e c t of va r i a t ions i n c r i t i c a l cont ro l parameters. I n addi t ion, simulated emergency conditions a re pract iced, and an t ic ipa ted emergencies t h a t might r e s u l t from malfunctions of the engine, i n e r t i a l platform, flow-direction sensor, rad io or radar , and s t a b i l i t y augmentation system a r e rehearsed. Fa i lures t h a t do not a f f e c t the f l i g h t a r e dis t inguished from those t h a t cause a l t e r a t i o n t o the f l i g h t plan. A n extreme f a i l u r e may d i c t a t e t h a t the f l i g h t plan be abandoned, with safe recovery of t he a i rp lane and p i l o t the only object ive.

Ground Control ler

Although the p i l o t i s i n complete cont ro l of t he f l i g h t , the ground moni- t o r ing s t a t i o n performs an important function i n support of the f l i g h t opera- t i o n . The monitor s t a t i o n i s equipped with displayed radar and selected channels of telemetered da ta . The ground con t ro l l e r s monitor subsystem opera- t i on , captive f l i g h t t rack , engine operation, f l i g h t t rack , s t a b i l i t y and con t ro l parameters, p i l o t ' s physiological fac tors , a i rp lane energy t o insure recovery, and, i n an emergency, d i r e c t a i r search and rescue.

Since a l l the f l i g h t s considered herein were made without on-board energy- management d isp lay f o r the p i l o t , the ground con t ro l l e r s furnished information t o the p i l o t as required. A s implif ied system, based on the energy-management foo tp r in t concept and cons is t ing of a family of curves f o r po in ts along the

projected track, was used to give the ground controller an indication of the range capability of the airplane at any time during the flight so that he could relay information to the pilot. The procedure has been satisfactory for all the flights; however, a ground-based real-time computation of energy has been mechanized and will be displayed as range capability for future flights.

RESULTS AND DISCUSSION

Two series of buildup flights were made. During the first, an altitude of 314,750 feet was reached with the basic airplane configuration. series was made with the ventral-off configuration, and an altitude of 354,200 feet, the maximum altitude obtained to date, was reached.

The second

The altitude and velocity of flights to high altitude considered in this paper are shown in fligure 5; detailed information concerning the flights is presented in tables I1 and 111. The boost portion of the flight to high altitude is analyzed in reference 11. As indicated in table 111, the pilot did not dways control to the desired altitude; however, during most of the flights he was able to control as accurately as the information displayed to him allowed. The boost-acceleration environment did not affect the pilot's performance.

High-Altitude Experience

During the flights to high altitude, aerodynamic controls were used until their effectiveness became low; the transition to reaction controls was then made. Below a dynamic pressure of 10 lb/sq ft, the reaction controls provided more control than could be obtained with aerodynamic controls (fig. 6). How- ever, the aerodynamic controls require large surface deflections which provide effective trimming moments during low-dynamic*-pressure flight but are not especially useful as damping devices.

Time histories of representative flights with reaction controls are pre- Summarized in the figures and in table I1 is sented in figures 7(a) to 7(h).

the experience with several types of control systems: acceleration command, rate command, and attitude command or hold operation for various piloting tasks. The tasks are shown as heavy lines. Table I1 also includes the pilot ratings of the reaction controls and piloting tasks and gives an estimate of the reaction control propellant used and the frequency of control utilization. Simulator studies (ref. 12) have shown that the effective use of reaction controls is a function of piloting experience and technique. This was particularly true with acceleration command reaction controls where the controlled quantity is two integrations removed from the immediate response to control. It is possible, therefore, that some of the data presented herein were, of necessity, obtained before the pilots had progressed to a steady performance level; however, in all cases, the pilots had sufficient practice on the simulator to be very familiar with the simulated reaction control piloting task.

Evaluation of acceleration command control.- Figure 7(a) presents data from a flight to a peak altitude of 2 l 7 , O O O feet in which acceleration command

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r eac t ion cont ro ls w e r e used. This w a s t he p i l o t ' s f i rs t f l i g h t a t low dynamic pressure. ca l led , i n i t i a l l y , f o r maintaining zero angle of a t t ack . A 5 O change i n angle of a t t a c k w a s requested a t maximum a l t i t u d e ; however, t he des i red angle w a s overshot. not achieved. The e f f e c t of t h e a i rp lane aerodynamics i s shown by the sinus- o i d a l o s c i l l a t i o n of the a i rp lane response i n both p i t c h and yaw. The d i s - turbance i n p i t c h i s probably the result of reac t ion controls , but the aerodynamic r e s to r ing moments of the a i rp lane help sus ta in the a i rp lane motions. N o con t ro l w a s used i n yaw; however, o s c i l l a t o r y motions are evident, probably i n i t i a t e d by t h e a i l e ron cont ro l . d i f f i c u l t than a t zero dynamic pressure. The aerodynamic cont ro ls used by the p i l o t and the SAS are a l s o included, inasmuch as during most of t he f l i g h t the aerodynamic cont ro ls w e r e almost as e f f e c t i v e as t h e reac t ion cont ro ls .

The minimum dynamic pressure w a s 3.4 lb / sq f t . The p i l o t i n g t a s k

Also, a 14" angle of a t t a c k f o r en t ry w a s requested but again w a s

Control a t low dynamic pressure w a s more

Evaluation of r a t e command control.- D a t a from a f l i g h t t o an a l t i t u d e of

Close con t ro l of 223,700 f e e t i n which rate-command reac t ion cont ro ls were used are shown i n f igu re 7 ( b ) . The minimum dynamic pressure w a s 2.9 lb / sq f t . bank angle i s indicated, a t t he expense of a high number of con t ro l inputs . The desired f l i g h t plan i n p i t c h w a s not followed closely, but t he requested en t ry angle of a t t ack w a s es tab l i shed and held. No p i l o t o r automatic con t ro l inputs were observed i n yaw, s ince the p i l o t w a s evident ly s a t i s f i e d with the a i rp lane heading and the r a t e s developed were l e s s than t h e system threshold. Tota l reac t ion con t ro l f u e l used w a s s ign i f i can t ly less than t h a t used during t h e previously described f l i g h t with the acce lera t ion command con t ro l system ( f i g . 7 ( a ) ) . r a t ed the con t ro l t a sks as general ly sa t i s f ac to ry .

Although the two f l i g h t s were made by d i f f e r e n t p i l o t s , both

Ef fec t of reac t ion damping.- Figure 7(c) presents da ta from a f l i g h t a t low dynamic pressure i n which the acce lera t ion command reac t ion cont ro ls were used together with reac t ion damping augmentation. For t h i s f l i g h t t he p i l o t w a s asked t o con t ro l p i t c h and bank angle within 8" of 0 t o obtain p i l o t - performance da ta appl icable t o a fu tu re stellar-photographic program with the X - 1 5 . Thus, t he p i l o t had a c l e a r l y defined object ive, and the cont ro l t a s k w a s successful ly accomplished within the des i red l i m i t s . This f l i g h t w a s made by a t h i r d X - 1 5 p i l o t who, a t the t i m e , had accumulated most of t he reac t ion con t ro l experience achieved during the program. Tota l cont ro l fue l used w a s h igher than with the rate command system but l ess than with the acce lera t ion command system. The p i l o t consfdered the e f fec t iveness of t he cont ro l system f o r the given con t ro l t a sk t o De exce l len t .

Comparison of manual and ~~~ automatic controls.- Figures 7(d) and 7(e) provide a comparison of cont ro l t a sks during f l i g h t s t o an a l t i t u d e of about 25O,OOO f e e t (design a l t i t u d e ) where the minimum dynamic pressure w a s less than 1 lb /sq f t . i ve cont ro l system w e r e used i n t h e f l i g h t depicted i n f igu re 7 ( d ) . I n i t i a l l y , angle-of-attack hold w a s used. Bank angle and heading w e r e cont ro l led manually as maximum a l t i t u d e w a s approached; t he hold modes were used during en t ry . Bank angle w a s cont ro l led prec ise ly but required ca re fu l monitoring. Angle of a t t a c k and p i t c h angle w e r e not c lose ly control led except p r i o r t o entry. Aerodynamic con t ro l inputs are a l s o evident since, w i t h blended controls ,

The rate command and hold reac t ion con t ro l f ea tu re s of the adapt-

aerodynamic con t ro l i s commanded with the same s t i c k movement t h a t commands reac t ion cont ro l .

Figure 7(e) shows a similar f l i g h t with manual con t ro l through t h e acce l - e r a t ion command system. The excursions i n t h e cont ro l led quan t i t i e s show t h a t t he a i rp lane motions were not cont ro l led as c lose ly by the p i l o t with the acce lera t ion command cont ro l system as the combined p i lo t - au top i lo t cont ro l led the f l i g h t of f i gu re 7(d) . A t maximum a l t i t u d e the p i l o t w a s asked t o bank t o approximately 30" t o evaluate the effect iveness of t he b a l l i s t i c con t ro l sys- t e m . The lack of reac t ion damping i s evidenced by the o s c i l l a t o r y motions. l a t i o n i n yaw.

The angle reached w a s 28". The p i l o t w a s not ab le t o damp the a i rp lane o s c i l -

-30" t r o l same t r o l with t h a t t r o l roll

I n both f l i g h t s the p i t c h reac t ion con t ro l inputs were predominantly up i n order t o maintain p i t c h a t t i t u d e and t o e s t a b l i s h en t ry angle of a t t ack . With the rate command system, the p i l o t w a s content t o allow slow excursions i n yaw; whereas, with the acce lera t ion command system, he attempted t o minimize the excursions i n yaw. With the rate command system, the bank angle w a s held within 5 " , but a t t he expense of many con t ro l inputs . With the acce lera t ion command system, the p i l o t w a s asked t o bank t o and s t a b i l i z e a t an angle of

and then r e tu rn t o 0 " . Both p i l o t s successful ly performed the bank con- t a sk assigned. With the two systems the con t ro l f u e l used w a s about t he f o r t he roll con t ro l task. For the yaw s t a b i l i z a t i o n task, much more con- f u e l w a s used with the acce lera t ion command cont ro ls . The cont ro l t a s k the more sophis t ica ted cont ro ls w a s r a t ed as subs t an t i a l ly improved over with the simpler acce lera t ion command system. Although t h e reac t ion con- inputs are shown t o be the same amplitude f o r p i tch , roll, and yaw, the reac t ion con t ro l t h r u s t i s lower and, so, consumes less con t ro l f u e l .

Presented i n figure 7 ( f ) are da ta from a f l i g h t t o an a l t i t u d e of 271,700 fee t with a minimum dynamic pressure of l ess than 1 lb/sq f t . command cont ro ls were used by the p i l o t t o con t ro l t he a i rp lane manually. The p i t c h and yaw modes i l l u s t r a t e loose but e f f ec t ive containment of t he motions with l i t t l e expenditure of cont ro l f u e l . The f l i g h t w a s flown by the most experienced X-15 p i l o t . inputs . It i s in t e re s t ing t o note t h a t t he p i l o t r a t ed the ro l l - con t ro l t a s k superior t o the p i t c h and yaw modes, which were cont ro l led less c lose ly . The t o t a l con t ro l f u e l used w a s much less than f o r some of t he lower-al t i tude f l i g h t s . The reduction i s a t t r i bu ted , f o r t he most pa r t , t o the p i l o t ' s g rea t e r experience i n using the reac t ion cont ro ls .

Rate

Bank angle w a s c lose ly control led with many cont ro l

F l igh t s above 3OO,OOO feet .- Near-maximum use w a s made of the automatic cont ro l f ea tu re s of the adaptive cont ro l system i n the f irst f l i g h t t o an a l t i - tude g rea t e r than 3OO,OOO f e e t ( f i g . 7 ( g ) ) . R o l l and heading hold were used during the e n t i r e f l i g h t , and bank angle w a s control led within close l i m i t s . Although a heading of 205" w a s specif ied, t he heading hold accepted by the p i l o t was about 4" less . Few cont ro l inputs w e r e required i n t h i s cont ro l mode. out, although a t the extreme a l t i t u d e the angle of a t t a c k w a s expected t o be unre l iab le because of the extremely low dynamic pressure. From maximum a l t i - tude t o the acquis i t ion of reent ry angle of a t tack , p i t ch -a t t i t ude hold w a s engaged and t h e system held p i t c h angle t o within 2" t o 3" of the des i red

An angle of a t t ack of 10" w a s requested during the f i n a l phase of climb-

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value. A s w a s expected with hold modes i n operation, t h e con t ro l system opera- t i o n w a s evaluated as sa t i s f ac to ry even i n yaw where some d r i f t w a s allowed. However, reac t ion con t ro l fue l consumption w a s r e l a t i v e l y high, espec ia l ly i n the p i t c h channel ( t a b l e 11).

A compromise between automatic hold and manual ra te command con t ro l ( f i g . 7 (h ) ) w a s used during a f l i g h t t o 347,800 fee t i n which the minimum dynamic pressure w a s ca lcu la ted t o be about 0.01 lb /sq f t . w a s used, and the system performance w a s ra ted sa t i s f ac to ry i n holding bank angle t o within 5O. accomplished manually by the p i l o t using the rate command con t ro l system. Average p i l o t i n g e r r o r i n p i t c h and en t ry angle of a t t a c k w a s about 5 O . More reac t ion con t ro l fue l w a s used i n the mode cont ro l led manually, but the p i l o t evaluated con t ro l i n a l l t h ree modes as exce l len t .

R o l l and yaw hold

The primary t a s k w a s con t ro l of t he p i t c h mode, which w a s

Summary of Reaction Control Experience

F l igh t experience with the X-15 a i rp lane has indicated t h a t the p i l o t uses reac t ion cont ro ls t o surpr i s ing ly high levels of dynamic pressure, where aerodynamic cont ro ls would be several t i m e s more e f f ec t ive than reac t ion cont ro ls . For example, reac t ion cont ro ls have been used a t dynamic pressures as high as 200 lb /sq f t on e x i t and 500 t o 600 lb / sq f t on reent ry . appears t o be pecul ia r t o the X-15 operation, i n which dynamic pressure de- creases a t a high r a t e during e x i t and bui lds up rap id ly during reent ry . Rates of change of dynamic pressure from about 30 t o 60 lb / sq f t / s e c have been observed during e n t r i e s from high a l t i t u d e s . During these c r i t i c a l con t ro l times, t he p i l o t used whatever con t ro l seemed most e f f e c t i v e t o accomplish the con t ro l t a sk . t he a i rp lane o s c i l l a t o r y motions during en t ry a t r e l a t i v e l y high dynamic pressure . These cont ro ls apply a pure moment with l e s s l a g and, so, do not exc i t e unwanted responses. The adaptive system reac t ion controls , by design, a l s o operate i n regions of rap id ly changing and high dynamic pressure, s ince these cont ro l regions a re c r i t i c a l for con t ro l l ab i l i t y , as i s discussed l a t e r .

This technique

The p i l o t s have used the reac t ion cont ro ls e f f e c t i v e l y t o damp

As noted previously, t a b l e I1 summarizes the f l i g h t experience obtained with reac t ion cont ro ls . Although t o t a l f u e l used does not necessar i ly r e f l e c t t he e f f ic iency or c o n t r o l l a b i l i t y of a system closed by a p i l o t , the t o t a l f u e l and con t ro l inputs a re shown and include experience a t a l l l e v e l s of dynamic pressure. For these f l i g h t s t he p i l o t s were thoroughly rehearsed on the f ixed- base simulator; nevertheless, these da t a must be considered t o be preliminary because of the l imi ted t o t a l experience i n these f l i g h t regimes.

Fuel consumption.- A s might be expected, t he reac t ion con t ro l f u e l consumption i n regions of high dynamic pressure w a s high and w a s not representa- t i v e of t h a t required t o s t a b i l i z e i n low-dynamic-pressure regions f o r which the reac t ion cont ro ls w e r e designed. I n order t o compare reac t ion f u e l consumption i n regions of low dynamic pressure, a t abu la t ion w a s made of t he use of t h e reac t ion cont ro ls i n regions of dynamic pressure l e s s than 5 lb / sq f t . Figure 8 ind ica t e s t h e t i m e ava i lab le f o r evaluat ion of t h e r eac t ion cont ro ls i n th i s range of low dynamic pressure and shows the maximum avai lab le cont ro l

9

time t o be of t he order of 3 minutes. It i s a l s o apparent t h a t f o r t he range of t he f l i g h t s t h e only va l id comparison i s on a rate-of-fuel-usage b a s i s r a t h e r than t o t a l f u e l consumed.

The reac t ion con t ro l fue l used per second f o r s t a b i l i z i n g t a sks i n f l i g h t regions with a dynamic pressure less than 3 l b / sq f t i s shown i n f igu re 9 f o r t h e fou r reac t ion con t ro l systems t e s t ed . Compared are f u e l rates f o r p i tch , r o l l , and yaw cont ro l . I n view of the pauci ty of data , the f u e l rates f o r p i t c h and roll con t ro l were comparable. concerned with c lose con t ro l i n yaw (heading), bu t indicated t h a t the yaw t a s k would be as d i f f i c u l t as p i t c h and roll i f t h e p i l o t chose t o cont ro l t h i s quant i ty c lose ly . This i s substant ia ted i n t h e figure by the experience shown f o r t he acce lera t ion command system. Fuel consumption of t he rate command cont ro ls and the automatic hold controls , which comprise most of t he data , w a s comparable. The l i m i t e d r e s u l t s obtained with the var ious systems obviate more de t a i l ed comparisons.

The p i l o t s appeared t o be the least

The reaction-system f u e l requirement f o r damping appeared t o be about one h a l f t h a t expended with manual cont ro l . The expenditure of f u e l f o r augmenta- t i o n appears t o be j u s t i f i e d when the superior p i l o t r a t i n g assigned the augmented cont ro ls i s considered. These data are discussed i n more d e t a i l l a t e r .

P i l o t performance.- Although the amount of r eac t ion cont ro l f u e l used may give some indica t ion of t he e f fec t iveness of t he system, a performance measure i s a l s o needed t o ind ica te the e f fec t iveness of t h e pilot-system i n accom- p l i sh ing the des i red cont ro l t a sks . The des i red a i rp lane a t t i t u d e s were ind i - cated on the example time h i s t o r i e s of t he h igh-a l t i tude f l i g h t s presented i n f igu re 7. The average absolute e r r o r during these f l i g h t s a t low dynamic pressure i s summarized i n f igu re 10. Performance i s summarized only f o r t h a t p a r t of the f l i g h t i n which the p i l o t w a s asked t o con t ro l t o a specif ied a t t i t u d e . From these meager data , it appears t h a t s l i g h t l y l e s s average e r r o r w a s achieved by the automatic hold mode of operation; however, acceptable cont ro l performance f o r t he X-15 high-a l t i tude missions w a s obtained with a l l systems. Larger average e r r o r s were evident i n the p i t c h control , as might be expected, s ince the s teep e x i t and en t ry angles required near-continuous compensation t o con t ro l p i t c h a t t i t u d e , simply as a r e s u l t of the f l i g h t t r a j ec to ry . R o l l e r r o r s were apparently c lose ly control led, inasmuch as they a re more e a s i l y detected by using cockpit d i sp lays and ex te rna l v i s u a l cues. Although the p i l o t s could con t ro l yaw angle c losely, they d id not choose t o cont ro l as c lose ly as the automatic system.

P i l o t ratings.- To summarize the f l i g h t experience with the reac t ion con t ro l systems, t h e p i l o t r a t i n g s of t he p i l o t i n g t a s k and cont ro l systems a re summarized i n f igu re 11. Although a l l t he p i l o t i n g t a sks with the cont ro l systems were r a t ed as sa t i s f ac to ry , t he con t ro l t a s k with the more sophis t i - cated systems ( ra te o r a t t i t u d e command) w a s r a t ed as s l i g h t l y improved over the simpler "accelerat ion command" system.

Considering the ove ra l l evaluat ion of t he reac t ion cont ro l systems tes ted , the simplest system--the acce lera t ion command reac t ion control--although satis- fac tory f o r the design goals of t he program, must be r a t ed somewhat less e f f e c t i v e than the o ther con t ro l systems. Fuel consumption could be high

10

depending on control technique, yet the control errors (fig. 10) were not significantly greater than with the other systems. eration command system, however, as slightly inferior to the other systems. Although the hold modes were appreciated for the roll and yaw controls, a preference for manual control of pitch--the primary mode of control--was expressed.

The pilots rated the accel-

Evaluation of Entry Flight Control

An airplane capable of a maximum Mach number of about 6.0 reenters the atmosphere from high altitude steeply at a large negative flight-path angle. This maneuver is challenging to the pilot, since it is flown at a relatively high angle of attack and under rapidly changing conditions of dynamic pressure and velocity, with the associated changes in aircraft stability and response. A time history of the most severe entry experienced to date (i.e., from the highest altitude) with the X-15 is shown in figure l2(a) in terms of dynamic pressure and dimensional aerodynamic derivatives. Figure 12(b) shows the lateral-directional and longitudinal undamped natural frequency and damping (unaugmented). sional derivative is primarily the result of the rapid change in dynamic pressure; however, the basic airplane characteristic motion, even at high dynamic pressure, is lightly damped. The rapid change in the stability and control sensitivity of the basic airplane with the light damping (unaugmented) was predicted from the flight simulator tests to be a severe control task, near the limit of the control capability of the pilot.

With Mach number nearly constant, the increase in the dimen-

With the augmented control system operating, the primary piloting task during entry is one of trimming the stabilizer to an angle that will result in the desired angle of attack for entry as dynamic pressure builds up. With this technique, the pilot makes corrections only to the airplane's attitude to limit the normal acceleration to the desired value during entry.

X-15 entry experience from high altitudes is summarized in table 111, which includes a comparison of the planned and actual flight parameters and pilot ratings of the entry control task. Representative entries from these flights to high altitude are presented in figure 13 to illustrate the X-15 entry characteristics and the piloting control task with the various control systems available in the X-15 airplane.

Figure l3(a) presents an entry from an altitude of 226,400 feet with the conventional controls (manual attitude command in pitch and yaw and roll rate command in roll) with damping augmentation. decreased range during the entry and increased directional stability, especially at high angles of attack. and dynamic pressure was 1360 lb/sq ft. Control was performed manually, using longitudinal trim to establish the angle of attack for entry.

Speed brakes were extended 20" for

Maximum entry normal acceleration was 3.6g

A similar entry from an altitude of 223,700 feet flown with the adaptive rate command control system is shown in figure l3(b). was about 4g normal and l.5g longitudinal.

Peak entry acceleration An angle of attack of approximately

11

20" w a s held u n t i l dynamic pressure and the r e s u l t i n g acce lera t ion buildup d i c t a t ed a reduction t o about 5 " . 1215 lb /sq f t .

Maximum ent ry dynamic pressure w a s

A comparison of t he two e n t r i e s ( f i g s . l3 (a) and l 3 ( b ) ) shows s i m i l a r c o n t r o l l a b i l i t y with t h e two types of cont ro l systems--the conventional cont ro l system with SAS and the adaptive cont ro l system. The reac t ion augmentation cont ro l system ( r e f . 9) w a s used f o r t he i n i t i a l p a r t of the en t ry from 226,400 f e e t ( f i g . l 3 ( a ) ) , and the adaptive con t ro l system, which provides r a t e command reac t ion control , w a s used during the e a r l y por t ion of the en t ry from 223,700 f e e t . Establ ishing en t ry angle of a t t a c k w a s a simple t a sk with e i t h e r cont ro l system. The lower ven t r a l had been removed f o r these f l i g h t s .

Figures l 3 ( c ) and l 3 ( d ) present data f o r e n t r i e s from the design a l t i t u d e of 250,000 f e e t with the conventional and the adaptive cont ro l systems. the en t ry of f igu re l 3 ( c ) , the a i rp lane w a s not equipped with reac t ion augmen- t a t ion , so the conventional acce lera t ion command cont ro ls were used during the ea r ly p a r t of the en t ry . The en t ry with the adaptive cont ro l system ( f i g . l 3 ( d ) ) w a s , f o r the most pa r t , made using the system hold modes. The ventral-on a i rp lane configuration w a s used f o r both f l i g h t s , and the speed brakes were extended 35" f o r the en t ry of f igu re l 3 ( c ) t o reduce the range. The l a t e r a l - d i r e c t i o n a l c o n t r o l l a b i l i t y with the adaptive cont ro l system, with the hold modes a, cp, and I) functioning, w a s superior t o the en t ry with the conventional cont ro ls with the lower-gain s t a b i l i t y augmentation system. With the low-gain damper system, a l a t e r a l - d i r e c t i o n a l o s c i l l a t i o n of l a r g e r amplitude i s evident during the i n i t i a l p a r t of the dynamic-pressure increase. How- ever, each p i l o t ra ted the cont ro l t a sk as sa t i s f ac to ry .

For

Other evaluat ions of the adaptive cont ro l system were obtained on the f l i g h t s shown i n f igu res l 3 (e ) and l 3 ( f ) . ( f i g . l 3 ( e ) ) w a s flown manually but with the r a t e command controls with the ventral-off configuration. using angle-of-attack, bank, and heading hold i n the ventral-on configuration. Speed brakes were extended 20" f o r the f l i g h t shown i n f igure l 3 ( e ) and 35" f o r the f l i g h t i n f igure l 3 ( f ) . en t ry ( f i g . l 3 ( e ) ) appeared t o be superior t o t h a t with the a t t i t u d e command o r hold mode. Although ventral-on and addi t iona l speed-brake extension increased the d i r ec t iona l s t a b i l i t y of the airplane, both f a c t o r s made the a i rp lane l e s s control lable i n r o l l . On the bas i s of these r e s u l t s , addi t iona l wind-tunnel t e s t s , simulator c o n t r o l l a b i l i t y s tudies , and system r e l i a b i l i t y considerations, it w a s decided t o f l y the remainder of the explorat ion program of the high- a l t i t u d e capabi l i ty of the X - l 5 with the a i rp lane equipped with the adaptive cont ro l system and with the lower ven t r a l removed.

The en t ry from 285,000 f e e t

The en t ry from 314,750 f e e t ( f i g . l 3 ( f ) ) w a s made

The c o n t r o l l a b i l i t y of the manually control led

The e n t r i e s from the two highest a l t i t u d e f l i g h t s t o da te a r e presented i n f igu res l3 (g ) and l 3 ( h ) . These f l i g h t s were flown manually i n p i t c h with the adaptive cont ro l system. The lower ven t r a l w a s removed, and the speed brakes were extended 20" f o r increased d i r e c t i o n a l s t a b i l i t y . Cont ro l lab i l i ty w a s ra ted very sa t i s f ac to ry i n both the low- and the high-dynamic-pressure regions. An ent ry angle of a t t ack of about 25" w a s held through acce lera t ion buildup t o about 58. There i s no evidence of the l a t e r a l - d i r e c t i o n a l o s c i l l a t i o n

12

cha rac t e r i s t i c s t h a t were found t o be objectionable during t h e e n t r i e s from 25O,OOO f e e t with the ventral-on configuration.

Summary of Entry Ekperience

Although t h e f l i g h t data t h a t have been obtained a re not su f f i c i en t t o determine the entry angle of a t t ack and normal accelerat ion required f o r recovery of the X - 1 5 from high a l t i t u d e , simulated f l i g h t da ta a re avai lable t o ind ica te these requirements ( f i g . 14) both f o r the clean a i rp lane from an a l t i - tude of 250,000 f e e t and the a i rp lane with speed brakes extended from a l t i t u d e s of 250,000 f e e t and 350,000 f e e t .

The X - 1 5 en t ry pul lout requirements, from the simulator, f o r the design

The simulator minimum veloci ty a t maximum

For ver i f i ca t ion , the peak entry accelerat ion and dynamic pres-

a l t i t u d e of 25O,OOO f e e t and a veloci ty a t maximum a l t i t u d e of 4450 f t / s ec a re shown i n f igures 14(a) and 14(b ) . a l t i t u d e w a s se lected t o be the average of the i l l u s t r a t e d f l i g h t s t o 250,000 f e e t . sure from f igures l3(c) and l 3 ( d ) are included, although the entry ve loc i t i e s were not i den t i ca l . Nevertheless, t he f l i g h t and s i m u l a t o r data, including the average angle of a t t ack during entry, agree.

Similar requirements f o r en t ry f r o m 350,000 f e e t with the speed brakes extended 20° are presented i n f igure 1 4 ( c ) . included and compares well with the predicted en t ry angle of a t tack, normal accelerat ion, and dynamic pressure.

The l imited f l i g h t experience i s

The performance and the con t ro l l ab i l i t y of the X-15 a i rp lane have, i n general, been closely predicted by the X - 1 5 complete six-degree-of-freedom f l i g h t simulator ( r e f . 13) . The simulation used the generalized equations of motion ( r e f . 14) referenced t o the airplane body axes.

Entry prof i les . - From these simulated f l i g h t s , it i s apparent t h a t the recovery t o l eve l - f l i gh t a l t i t u d e i s a strong function of entry angle of a t tack and, t o a l e s s e r extent, the normal accelerat ion held f o r pu l lou t . The combination of angle of a t t ack and pul lout accelerat ion determines the peak value of dynamic pressure during entry. angle of a t t ack and v e r t i c a l accelerat ion i s required t o avoid pul l ing up instead of es tab l i sh ing a control led g l ide . The entry experience obtained during the X - 1 5 e n t r i e s from high a l t i t u d e i s summarized i n f igu res 15 t o 18.

A s dynamic pressure bui lds up, a reduction i n

The time accumulated a t elevated g during e n t r i e s ( f i g . 15) has ranged from minimal time a t 5g t o nearly 10 minutes a t l .5g .

Summarized i n f igures 16(a), 16(b) , and 1 6 ( ~ ) , respect ively, a r e the ranges of Mach number, dynamic pressure, and angle of a t t ack encountered during the en t ry and terminal g l ide from the high-al t i tude f l i g h t s . Since f l i g h t s t o higher a l t i t u d e require more t i m e than those t o lower a l t i t u d e , the time base of t he en t ry and ranging da ta presented w a s sh i f ted t o a l i n e the increase i n en t ry dynamic pressure. Maximum Mach numbers ( f i g . 16(a) ) during en t ry ranged from 4.25 t o 5.4. glide-to-base port ions of t he f l i g h t and resu l ted i n increased range.

As planned, subsonic Mach numbers were used during the

1 3

I

1

The maximum dynamic pressure obtained during en t ry i s dependent on the maximum a l t i t u d e , angle of a t tack, and the p i lo t ing technique used during entry. Figure 16(b) shows the range of en t ry dynamic pressure t o be f rom 800 lb/sq f t t o 1800 lb /sq f t . t o base w a s lower and varied from 200 lb/sq f t t o 400 lb/sq f t .

The dynamic pressure during the g l ide t o re turn

During t h e aerodynamic port ions of f l i g h t s t o high a l t i t u d e , the angle of a t t ack varied over a range of about 0" t o 28" during en t ry ( f i g . 1 6 ( ~ ) ) . an angle of a t t ack of about 10" r e s u l t s i n near maximum l i f t - t o -d rag r a t i o (approximately 2 .5) , it appears t h a t the X-15 f l i g h t s were planned and flown without the need f o r f ly ing maximum L/D. generally l e s s than t h a t f o r maximum L/D.

Since

Glide angles of a t t ack flown were

P i l o t evaluations.- P i l o t ra t ings of the en t ry cont ro l t a sk with the various cont ro l systems are presented i n f igu res l 7 ( a ) , (b) , and ( e ) as a function of maximum dynamic pressure, average angle of a t tack, and maximum normal accelerat ion. ~n only one instance ( f i g . 1 7 ( ~ ) ) w a s an en t ry cont ro l t a sk ra ted poorer than sa t i s fac tory , and no degradation i n p i l o t r a t i n g i s indicated over t he range of entry parameters obtained. Also, no preference f o r a pa r t i cu la r aerodynamic control system i s indicated. Although the yaw hold mode on one f l i g h t was rated by the p i l o t t o be l e s s s a t i s f ac to ry than manually f ly ing the airplane, adjustments t o the system's dead band have resu l ted i n a system considered t o be completely sa t i s f ac to ry .

Control of range.- Maximum range a t ta ined during en t ry f l i g h t w a s obtained during the f l i g h t t o an a l t i t u d e of 354,200 f e e t ( f i g . 18). the speed brakes were extended 20" f o r increased d i r ec t iona l s t a b i l i t y a t high angles of a t tack . nominally an a l t i t u d e of 3O,OOO f e e t and a Mach number of 0.8, w a s approximately 270 nau t i ca l miles. The duration of en t ry w a s only about 3 minutes.

For t h i s entry,

The t o t a l range t o landing pa t t e rn high key, which i s

The l i f t - t o - d r a g r a t i o f o r the X-15 a i rp lane a t supersonic speeds i s presented i n f igure 19 t o indicate the range of l i f t - d r a g r a t i o avai lable t o the p i l o t through cont ro l of the speed brakes. Data a re shown f o r an angle of a t t ack of l oo , which gives approximately maximum L/D. Additional values of L/D below the maximum value can be obtained by the p i l o t through control of angle of a t tack ; however, extensive use of t h i s cont ro l would r e s u l t i n unwanted excursions i n the f l i g h t path. Although angle of a t t ack i s normally used f o r long-range control, the speed brakes may be used f o r e i t h e r long- o r short-range control, such as during approach and landing.

This f l e x i b i l i t y i n control l ing range during en t ry by the use of speed brakes and turning f l i g h t i s i l l u s t r a t e d i n f igure 20, which compares an entry from an a l t i t u d e of 354,200 f e e t with various simulated f l i g h t da ta f o r the same mission. During the f l i g h t shown, the p i l o t used speed brakes intermit- t e n t l y a t h i s own d iscre t ion . A range of about 200 miles i s avai lable t o the p i l o t from maximum a l t i t u d e a t a veloci ty of 4300 f t / s ec , but with speed brakes the forward range can be reduced t o about 140 m i l e s . If a 2g tu rn i s used for a 180" change of heading, the forward range can be minimized t o about 85 miles. P i l o t comments indicate t h a t the landing s i t e can be observed f rom the highest a l t i t u d e reached t o date and f rom a range of 160 miles; however, correct ive cont ro l toward the s i te cannot be made u n t i l the en t ry i s nearly completed.

14

Yet, through careful flight planning, only one incident has occurred that resulted in marginal recovery at the selected landing site. On this occasion, the pilot nearly overflew the landing site by "bouncing" on entry.

The recovery and ranging plan for flights from high altitude to a high key position of h = 30,000 feet and M = 0.8 over the Edwards landing site is illustrated in figure 21. The flights to high altitude with a minimum velocity at maximum altitude of 4000 ft/sec to 3000 ft/sec are planned with 50 to 100 miles of excess range to insure landing at the desired landing site. Com- pared to the actual flight experience (crosshatched region) is the minimum recovery range for Edwards landings. Range for entry from altitudes greater than 200,000 feet is planned to be greater than 180 nautical miles; flights to 100,000 feet require a range of l5O nautical miles or less.

Additional information concerning the pilot's control of range during recovery and landing of the X-15 airplane is summarized in reference 15.

CONCLUSIONS

Piloted flights to high altitudes have been successfully accomplished with several different control systems in two X-15 airplane configurations. The performance of the pilot and his evaluation of the control tasks during the low-dynamic-pressure and entry portions of high-altitude flight have been analyzed to indicate the effectiveness of the control used and the severity of the control task. These flight tests led to the following conclusions:

1. The pilot's performance during flight at low dynamic pressure was similar with all of the reaction control systems tested--angular acceleration, angular acceleration with damping, and rate command--and was comparable to the performance of the automatic hold systems. The pilots evaluated the reaction control task with all of the systems as satisfactory; however, they preferred the automatic hold modes in roll and yaw and manual control in pitch. Reaction control fuel usage was generally higher with the less sophisticated acceleration command control system.

2. For the entry control task from altitudes of 200,000 to 354,200 feet all of the aerodynamic controls--attitude command with damping, rate command, and attitude hold--were rated satisfactory by the pilots. The entry acceleration environment did not affect the pilots' performance.

3. With the preflight procedures established for the X-15 program, no range control problems have been experienced during terminal glide to the landing site. The pilots have been able to judge and control excess range by modulating angle of attack and using speed brakes to insure landing at the desired landing site.

Flight Research Center, National Aeronautics and Space Administration,

Edwards, Calif., December 23, 1965.

11111ll11l11111l11111lll11IIl I I I I I 1 I1 I I I1 I1

SYMBOLS

A l l quantities are referenced to an airplane body-axis system.

ax longitudinal acceleration of airplane center of gravity, g units

transverse acceleration of airplane center of gravity, g units aY

a, vertical acceleration of airplane center of gravity, g units

b wing span, feet

Lift force lift coefficient, is CL

Rolling moment GSb

rolling-moment coefficient, C l

3% cb = as

Pitching moment pitching-moment coefficient, Cm [SC

16

Cn

CY

C

D

h

=Y

Yawing moment CSb

yawing-moment coef f ic ien t ,

Side force side-force coe f f i c i en t , cis

wing mean aerodynamic chord, f e e t

drag force, pounds

acce lera t ion due t o gravi ty , feet/second2

geometric a l t i t u d e , f e e t

moment of i n e r t i a re fer red t o roll body ax is , slug-foot2

moment of i n e r t i a re fer red t o p i t c h body ax is , slug-foot2

1, moment of i n e r t i a re fer red t o yaw body ax is , slug-foot2

L l i f t force, pounds

'sb2 C , per second %=q 2p

= 3 cis C h , per second

Lp = qSb Czp, per second 2

Lg = -Sb Cz8, per second 2 IX

M

% =

% =

q+

M6 =

m

Nr =

Np =

N6 =

P

4

q

r

S

t

v

-

Mach number

2 qsc Cm , per second 2v IY 9

qsc %, per second 2 IY

2 qsc 2v IY C%, per second

;is. 2 I Y 8

Cm , per second

m a s s , s lugs

-Sb2 Cnr, per second

-Sb Cn , per second 2 Iz

$33 Cn6, per second 2 rz r o l l i n g ve loc i ty , radian s/second

p i tch ing ve loc i ty , rad ians/se c ond

dynamic pressure, pounds/square foot

yawing ve loc i ty , radians / se c ond

wing area, square f e e t

time, seconds

t r u e veloci ty , feet/second

YB = Cyp, per second

a angle of a t tack , degrees

B s i d e s l i p angle, degrees

Y f l i gh t -pa th angle, degrees

D incremental change

18

. . - . . .. . -. .. . ,

I - ~

6

c 8

rp

3

cu

Sub s c r i p t s :

a

e

h

max

min

0

r

V

e

rp

3

cont ro l def lec t ion , degrees

damping r a t i o

p i t c h angle, degrees

bank angle, degrees

heading angle, degrees

undamped n a t u r a l frequency, radians/second

a i l e ron

en t ry

hor izonta l s t a b i l i z e r

maximum

minimum

i n i t i a l

recovery t o l e v e l f l i g h t

v e r t i c a l s t a b i l i z e r

p i t c h reac t ion cont ro l

roll reac t ion cont ro l

yaw reac t ion cont ro l

A dot over a symbol ind ica tes a der iva t ive with respect t o time.

I 1l11l111111l1111ll I1 I I1 I II

REFEmNCES

1. Holleman, Euclid C.; and Reiser t , Donald: Cont ro l lab i l i ty of t he X-15 Research Airplane With Interim Engines During High-Altitude F l ights . NASA TM X-514, 1961.

2 . Yancey, Roxanah B.: t i v e s of the X-15 Research Airplane t o a Mach Number of 6.02 and an Angle of Attack of 25".

F l igh t Measurements of S t a b i l i t y and Control Deriva-

NASA TN D-2532, 1964.

3. Taylor, Lawrence W., Jr.: Analysis of a Pilot-Airplane Latera l I n s t a b i l i t y Experienced With the X-15 Airplane. NASA TN D-1059, 1961.

4. Tremant, Robert A.: Operational Experiences and Charac te r i s t ics of the X - 1 5 F l igh t Control System. NASA TN D-1402, 1962.

5. Taylor, Lawrence W., Jr.; and Merrick, George B.: X-15 Airplane S t a b i l i t y Augmentation System. NASA TN D-1157, 1962.

6. Ebskovich, Boris; Cole, George H.; and Mellen, David L.: Advanced Fl ight Vehicle Self-Adaptive F l igh t Control System. Par t 1 - Study. WADD Tech. Rep. 60-651 (Contract No. AF33(616)-6610), Wright Air Dev. Div., U.S. A i r Force, Sept. 30, 1960.

7. Lindahl, J. H.; McGuire, W. M.; and Reed, M. W.: Advanced Fl ight Vehicle Self-Adaptive F l igh t Control System. Par t V. Acceptance F l ight Tests. WADD-Tech. Doc. Rep. 60-651 (Contract No. AF33(616)-6610), Aero. Systems Div., U.S. A i r Force, May 1963.

8. Lindahl, J.; McGuire, W.; Reed, M.: Advanced Fl ight Vehicle Adaptive F l igh t Control System. Par t V I I : F ina l Report on Study, Development, and Test of the MH-96 System for the X - 1 5 . (Contract No. AF33(616)-6610), Res. and Tech. Div., U.S. A i r Force, Dee. 1963.

WADD Tech. Rep. 60-651

9. Ja rv is , Calvin R.; and Lock, Wilton P.: Operational Experience With the X - 1 5 Reaction Control and Reaction Augmentation Systems. NASA TN D-2864, 1965

10. Hoey, Robert G.; and Day, Richard E.: Mission Planning and Operational Procedures for t he X-15 Airplane. NASA TN D-1159, 1962.

ll. Holleman, Euclid C.: P i lo t ing Performance During the Boost of the X - 1 5 Airplane t o High Alt i tude. NASA TN D-2289, 1964.

l2. S t i l l w e l l , Wendell H.; and Drake, Hubert M.: Simulator Studies of J e t Reaction Controls f o r Use a t High Alt i tude. NACA RM H58G18a, 1958.

13. Holleman, Euclid C.; and Wilson, Warren S.: Flight-Simulator Requirements f o r High-Performance Aircraf t Based on X-15 Experience. No. 6 3 - ~ ~ ~ ~ - 8 1 , ASME, 1963.

Paper

20

14. Cooper, N. R.: X-15 Fl ight Simulation Program. ARS-IAS Paper No. 61-194-1888, ARS-IAS, June 1961.

15. Hoey, Robert G.: Horizontal Landing Techniques for Hypersonic Vehicles. AGARD Rep. 428, Jan. 1963.

21

TABLE I.- TYPICAL x-15 FLIGHT PLAN

X-15 F l i g h t Request

F l igh t N o . : 3-20-31 Scheduled date: June 27, 1963

P i l o t : Major Robert Rushworth

Purpose: Alt i tude buildup, evaluate horizon scanner, in f ra red experiment, and u l t r a - v i o l e t experiment

Launch: Delamar Lake on magnetic heading of 214" with adaptive f l i g h t control system on "adaptive" damper but with roll hold on, reac t ion cont ro ls "auto, 'I both b a l l i s t i c cont ro l systems "on, 'I and heading vern ier switch on "standby."

Item

1

2

3

4

5 6

7

8

9

10

11

Time , see

0

15 29

79

86 170

243

282

290

h, f t

45 , 000

45 , 000 50 , 000

150 , 000

180 , ooo 2 78,000

188 , ooo

85 , ooo

76 , ooo

V, f t / s e c

790

1400

2900

5100

5000

4250

5000

4600

3900

- 9,

lb / sq f t

145

440

5 30

60

8 0.2

10

600

760

Launch, l i g h t engine, increase t o 100-percent t h r u s t , and r o t a t e the a i rp lane u n t i l 2g i s a t t a i n e d .

Maintain 2g u n t i l

A t 8 = 42", engage e hold, use

Shut down the engine a t

0 = 42".

vern ier t r i m t o maintain 0 = 42".

V = 5100 f t / s e c 0 hold.

and disengage

Switch heading vernier t o A$.

A t maximum a l t i t u d e extend the speed brakes t o 20". R o l l i n t o 30" l e f t bank and re lease the controls . Maintain e = -5" u n t i l Q: = 23".

A t M = 23" turn roll hold off and switch heading vernier t o "standby." Maintain Q: = 23' u n t i l normal accelerat ion = 5g.

i = -700 f t / s e c . Madimum entry S = 760 lb/sq f t .

A t 6 = -700 f t / s e c push over t o Q: = 3', extend speed brakes t o 35", and vector t o high key. Turn reac t ion cont ro l switch and engine master switch "off . I t

Maintain az = 5 g u n t i l

A t high key use speed brakes a s required. Pressurize the tanks a t l 7 , O O O f e e t and check f l a p and "squat" c i r c u i t breakers i n .

After landing s l i d e out, before auxiliary-power-unit shutdown, cycle f l a p s , s e t s t a b i l i z e r t r i m t o zero, and turn a l l data "off ."

22

I

TABLE 11.- SUMMRW OF FLIGHTS WITH RJXCl'ION "TROIS

B 246,700 Rate camand m h h Hold a = 8', 'p = Oo, f = 205' 1.0 1.0 a, 9, 'P, h u n t i l h = 200,000 fee t , then t hold maintain e = 0' manually. A t

h-, engage a hold a t 8 = 0'. A t a = ZOO, engage a hold.

A 247,000 Acceleration m m m Maintain a = 10'. A t hPx, 3.0 1.0 ~ camand ' bank V = -30'. A t h = lB0,WO

fee t , es tab l i sh a = 18" f o r I entry.

7, ,- Reactim-control inputs I Bal l i s t ic control Reaction-control fue l used, l b %tal

p i l o t ra t ing ' reaction - control ' - Q ! * - Pi lo t ing task pilDt % Reaction-control

Bystems --- e P i 9 B ? lb D m Right Le f t Right l e f t fuel. l b UD D O ~ Riaht l e f t Rinht l e f t

1.0 18.5 0.8 6.7 8.0 0.4

3.0 17.7 5.5 5.0 6.4 14.2

A 193,600 Rate command ma m m Roll t o 30" r i g h t bank, maintain 1.5 1.5 1.0 55.6 1.8 0.3 5.1 10.0 1.9 '74.7 . 62 4 4 46 20 4 0 hold hb for 30 second6. kt &, engage

a hold a t a = 0 . A 209,400 Rate cOnmYUld m h h With 'p and f hold on, main- 1.0 1.0 1.2 1.0 2.0 2.3 3.3 0 0.2 8.8 3 ll 19 37 0 1

9, $ hold t a i n a = 6'. A t &, r o l l i n t o 30" l e f t bank and release. Maintain e = 0' u n t i l a = 20'. Maintain a = zoo.

C

B

A

A

B 217,wO Acceleration m m m Maintain a = B = 0' t o +. 4.0 2.0 cnmnand A t h, ro ta te t o a = 5 and

hold, then t o a = 0'. Maintain a = 14' far entry.

bax, maintain a = oD u n t i l a = 20 , &intain a = 20' for entry.

C 223,700 Rate command m m m Maintain a = 5' manually. A t 3.0 1.5

A 226,400 Acceleration m m m Maintain a u n t i l a = O', then 1.0 1.0 cOmmand 9 = B = 0' A t h-, main- damping t a i n 9 = 0 u n t i l a = 20'.

285,000 Rate command m h h With 0, $ hold an, manually 2.5 3.5 1.5 9.8 3.3 5.3 6.5 1.2 9, t bold con t ro l a . A t h-, maintain

a = -5" u n t i l a = 23'. Main- tain a = 23', q, * hold off for entry.

314,750 Rate command m h h With 'p, $ hold engaged, main- 2.5 1.0 1.5 35.3 3.1 8.2 U.3 0 a, 9, B, h t a i n a = 10' manuauy. A t i hold b, engage a hold a t 8 = 0'.

A t a = 23", engage a hold.

347,800 Rate command m h h 0, $ hold on. A t k x , main- 1.0 1.0 1.0 23.7 12.3 7.0 8.6 14.5 B, i hold t a i n a = -20' u n t i l a = 23".

Maintain a = 23" for entry. 9, $ hold off during entry.

354,200 Rate c o m d m h h 0, $ hold on. A t h = j30,OOO 1.2 1.0 1.2 17.5 14.0 6.3 3.4 5.1 9, i hold f e e t , push over t o a = O D . Roll

45' t o -45'. A t h-, maintain level f l i g h t and a = -20'. A t h = 220,000 feet, es tab l i sh a = 26' for entry v i t h q, $ hold off.

2.0 ----

2.0 8.5

1.1 30

~ A 1 271,700 1 Rate conrmand ~ m 1 m 1 m 1 Maintain a = 6'. A t h-, 1 3.0 11.0 1 3.0 1 7.1 1 9.7 110.3 1 ll.l 1 2.3 maintain a = 0' u n t i l a = 23', then maintain a = 23'.

---- 84.0 17 51 IJ 14 o o

0 24.0 17 20 38 52 0 0

14.9 72.0 80 42 16 38 ll 40

53 75 1 4

Iu W

9 hold, adapt ivea

adaptive

(8,6,81

Manual

ManualsAs

Manual adaptive

220,000

206,000

200,000

220,OW

Menual adaptive

Manual adaptive

a, c, t hold, Bdaptiw

Manual

250,OOC

278,wC

282,mc

j15,OOC

TABLE 111.- SUMMARY OF ENTRY DATA

- il0t

- A

A

B

C

A

B

A

A

C

B

A

A

-

confi - sntral

ration - Speed brakes

Planned - h, ft (level flight)

Actual - $ax> b/aq ft - 9%

844

1,436

1,215

1,360

1,086

1,020

1,889

1,205

1,186

1,362

1.207

Pilot rating I -

le8 - !O

!O

14

'0

?O

'0

18

23

23

23

23

26

- a% J

g - _ _ _ 4.0

4.0

4.0

4.0

5.5

5.0

5.0

5.0

5.0

5.0

5.2

- i n a x ]

b/sq ft

- Ye 1

de6 - -17

-28

-26

-28

-26

-28

-29

-25

- 35

-35

- 36

Control h, I ft

h, ft

:light)

57,000

:level

-

72,003

51, W O

j1,WO

j4,OOO

12,000

63,000

49,000

56,000

65,000

66,000

68.000

- c - 1

1

2

1

1

1

1

3

1

1

- Ugh - ?4

19

13

19.5

20

19

20.5

27

26

22.5

26

26.5

- igh

3.3

-

3.7

3.7

4

3.6

4.4

5

3.6

4.2

5.1

4.8

5.2

verage - 10

12

12

18.5

18

16

17

23

20

20

23

25

verage

2.7

-

3.2

3.4

3

3.2

3.5

4

3.2

3.8

4.9

4.3

4.5

on

Off

on

O f f

O f f

On

on

O f f

Off

on

Of f

Off

Extended

Extended

Extended

Extended

Extended

Retracted

Extended

Extended

Extended

Extended

Extended

ExtendeC

9.0

5.0

7.0

6.0

6.0

b0.75

b1.0

b0.6

b0.2

b0.4

b0.l

bo.1

193,600

209,400

217,000

223,700

226,400

246,700

247,000

271,700

285,000

314,750

347,800

354,200

-

10.6

4.0

3.4

2.9

2.0

b0.6

b0.g

b0.2

b0.2

b0.03

b0.01

bo.01

80,000

72,000

84,000

84,000

80,000

70,OCC

82,OCo

76,000

80,000

73,M)O

70,000

550

1,000

6 30

6 30

650

1,150

625

760

625

1,000

1,200

3

1.5

Manual SAS 220,000 (8,4,8)

1

1.5

1.:

1.5

4

--- --- ~ 1.5 1.5 1.5 2 4,070 ' 5,000

4,550 5,515

4,270 5,480 1 1 1 1.5 1

1.5 1 1.5 1 1

1

1 -37 4,310 5,510

,ll augmentatlon inadvertently turned off durlng entry. st estimate.

Figure 1.- X-15 airplane.

I 49.5

Lower movable ventral

Figure 2.- Three-view drawing of the X-15 a i rp lane . A l l dimensions i n f e e t .

26

ri! 4

E-10542 Figure 3 .- X-15 cockpi t .

6000

4000

V, ft/sec

2000

0

300

200

h, ft

loo

0

I 1600

I I I I

t, sec

Figure 4.- Representative f l i g h t t o high a l t i t u d e .

28

h, ft

3 6 0 ~ 1 0 ~

320

280

240

200

160

120

ao

40

-

-

-

- Burnout

-

-

- Lau

-

I I I I I I I I I I I I I I I I I 0 0 0 0 0 0 0 0 0 0 0 0 1 2 3 4 5 Ax103

V, ft/sec

Figure 5.- X - 1 5 altitude flights considered during this investigation.

1.2

.8

Pitching angular

acceleration, 1 /sec2

.4

0

Rolling angular


Yawing angular


Control

~ Aerodynamic

_-- Reaction

.A

Figure 6.- Comparison of the aerodynamic and reaction control effectiveness. 6h trimmed at -25' f o r CXe = 20'.

- Planned 100 -

q r Ib/sq ft

0 I I I

s", 0

UP

Down se

-- ~

I 1 I I I I I

-40 I I

-40 ' 1 I I I I I 1

Right I 0 I i n n I n i I I sv Left I I u I I I I u 11 I I I I I I I

10 I 1

I I I I I

Right -1 sdJ

Left I I I I I I I

(a) Acceleration command controls manually operated, hmax = 217,000 ft, ?&in = 3.4 lb/Sq ft.

Figure 7.- Time histories of flights at very low dynamic pressure showing planned and actual values of airplane attitudes.

I I I ! 1 1 I 1

20

Sa, 0

des

L. 1 I -20

I I I I I 210 230 250 270 290 310 330 350

t, sec

(b) Rate command c o n t r o l s manual ly ope ra t ed , hmax = 223,700 f t , qmin = 2.9 lb/sq ft . Figure 7 .- Continued.

32

-Planned - q,

Ib/sq ft 0

40

UP

" Down

1 I I --

I I n I III m ~IIIII I II I I I 11111 I 1111 IIII

11111 \ Ifol"u II

I I I I I I I I

- 4 0 I 1

40 I I Pt

des 0

-40

- I _ - W

L y = O o ( + 8 O )

1 1 1 I I I I

Right I I I I II I I 111 69 Left 1 ni I I 1 I I II II 11111 II I I I I I II

I I I I

2o I 1 -20 ' I 1 I

220 I- I I I I

* -- ---- 2o r r - L~ = 214'

1 - I

I I 1 I I

200 220 240 260 280 300 320 340 360 t, sec

( c ) Acceleration command cont ro ls with r a t e damping manually operated, hmax = 226,400 f t , &in = 2.0 lb/Sq ft-

Figure 7.- Continued.

33

1l1111111111IlI 1111l I I Ill1 I I1

UP

Down so

} Planned - Manual --- Hold

100 - q r

Ib/sq ft 0 I I I I

I n o I n i i t n n n n i n n i i n n n n l I I I I I I I

0 a = 8O

40

0- \

L, = 00

I I I I I I I

0

Right

’+ Left

I I I I I I I I ! I

Right

le f t II I i n n i UI I I I iiiini nionti iinmn nnn linin1 I 1111IIIUIII I 1 0 1 Y u

sw 0

2 0 I I

-

I I I I I I I I

2 2 0 I I

- 2 0 ’

I I I I I I I I 180 ‘

(d) Rate command and a t t i t u d e command &in Y 0.6

Figure 7.-

(hold) controls , hmax = 246,700 f t , ib / sq f t .

Continued.

34

2o r

a, e, d eg

" P I I I I I I ni I I I 0 1 n in i Down

se I I I I I I I I

- 4 0 I I

Right I I II 1 n IIII 111 I I u [r n l MI I Left I-M I l l I I1 I

$t

d f g 1 deg

-20

20 I I

-20 ' I I I I I I I 1 I

I 180

I 160 ' I I I I I I I I I

n in I I I n Right l i r i i I IMI n 111 inti I I U I I I I I Left I

10 I 1 01 II~ ----- -

-10 I I I I I I I I

200 220 240 260 280 300 320 340 360 380 t, set

( e ) Acceleration command cont ro ls manually operated, hmax = 247,000 f t , &in 0.9 lb/Sq f t .

Figure 7 .- Continued.

35

IIIIIIIIII

I

- Planned - ql

Ib/sq ft

UP

Down 68 I

I II 011 I I I 1 I

40

4 0

.d

I r Q = 0" - oL- - - -

Right

'JI Left

'Q

n I I I I I

I I I I I I 1 I I I I

b 0

I I I 1 I 1 I 1 1 -40 ' 1

Ab -- I - - I

Right d " i n q Left

I I I 1 - ! 1 1 I -20 I 230

210

I I I I 1 . 1 I I I 190 I

( f ) Rate command controls manually operated, h m a = 271,700 ft, cmin 0.2 lb/sq f t .


Manual ) Planned Hold -

I I I I I I

- q,

Ib/sq ft 0

I Se UP b i n i m i I I I I I I ni I I p, I I II I I iiiniiiunn~ 1 1 t

Down [ I I I 1 I I I I I

Qi

-cy- 00 0- --

I

I I I I I I I I I 1

I I I I I I I I I -10 I 310 I I

Right I I I I I ' J I Left I I I I I I I I I I

l o I des -10 O r - I

230 2 5 0 270 2 9 0 310 330 350 370 390 410 430 t, sec

(g) Rate command and a t t i t u d e command (hold) cont ro ls , hmax = 314,730 ft, &in 0.03 lb/sq f t .


37

I v

4 0

0

- Manual --- Hold } Planned

I I l o o h - 1 I I I

- q,

Ib/sq ft 0

_______---__--_a__-------- -- -L,=23 -.-., - Q = - f O o

Q, e, des

4 0

0 - 9 des

-

rp = 0" v "

O /

160 deg

UP 11 I I I I W R I n W1

I I Down 'I se

-

I in I B n n u innnon I u II 1 I UI Y 1 v u - 1 1 1 1 I I I

- 4 0 sh,

des I

- 4 0 1 I I I I I I 1 I I I I I

I -10 1:E-~7-7- 1 1 - 1 I - I 1 1

s",

2 0 0 2 2 0 2 4 0 2 6 0 2 8 0 300 3 2 0 3 4 0 3 6 0 3 8 0 4 0 0 4 2 0 4 4 0 t. sec

(h) Rate command and a t t i t u d e command (hold) controls , hmax = 347,800 f t , &in 0.01 lb/sq f t .

Figure 7.- Concluded.

38

32C

28C

b"f ft

240

2 0 0

160 I I I I

t at 4 < 5 Ib/sq ft, s e c

40 8 0 120 160 2 0

Figure 8.- Control evaluation time available at a dynamic pressure less than 5 lb/sq ft. Velocity at apogee approximately 4,700 ft/sec.

39

UP .2

.1

Fuel usage rate,

Ib/sec

0

.1

Acceleration command

-

Acceleration Attitude command command

with damping Rate command (hold)

Right .1

Fuel usage

Ib/sec

U

Pitch

Right .1

Fuel usage

Ib/sec

Figure 9.- Reaction control

u Roll

u - ++€I- - - + Yaw

f u e l used, with t h e various cont ro l systems, per second of operation a t a dynamic pressure l e s s than 5 lb/sq f t . S t ab i l i z ing con t ro l t a sk .

40

Average absolute

error, d eg

Average absolute

error, deg

Average absolute

error, deg

20

10

0 -

'

Acceleration Attitude - @c;m,nd Acceleration , ~~ , witida1;ing command Rat;o,;;d command (ho.1

-

Pitch 0 0 08 o -

0 0 0

0 80 00 Yaw 0

Figure 10.- Summary of reaction control performance (attitude error) with the various control systems for the X-15 control tasks.

41

Satisfactory

Unsatisfactory

Unacceptable

Satisfactory

Unsatisfactory

Unacceptable

Satisfactory

Unsatisfactory

Unacceptable

Acceleration

Acceleration command Rate

command with damping command

0 e - - 0

0

Attitude

command

(hold)

0

0 0

0 8 0 0

0 0

0 0 0 0

Figure 11.- P i l o t r a t ing of the X-15 low-dynamic-pressure control t a s k with the various control systems.

42

32

24 a,

dag 16

8

1200

800 - q ,

Ib/sq ft

400

0 1

- -- l 2 I

La

LR

La V

N 8 V

-1 6 /- M a - - . . . . --

Ma Mat Ma

0 I

.4 r -.4

LP

- lo 0 1 8o I 1

,’ L8v

401--diLl 0 --- 8 1 I

N a p j 0

4 c ,

0 410 420 430 440 450 460 470 480

t, sac

(a) Dimensional stability and control parameters.

Figure 12.- Variation of the X-15 aerodynamic parameters during entry from an altitude of 354,200 feet. open 20°.

Lower movable ventral off; speed brakes

43

0,

rad/sec

.08

.04 -

0 41 0 420 430 440 4 5 0 460 470 480

t, sec

(b ) Pi tch and yaw undamped na tura l frequency and damping ratio.


44

1600 -

1200 -

8 0 0 - h, ft

4 0 0 -

0 -

- 4 0

32OX1O3 I

81 I

2 ' I I I 1 I I 1 I I I

- 4 0 7

-. ---/' 'Le __-- --------- ''$ ..-,---. #/--/ /.

4 0 l

3 2 0 3 3 0 3 4 0 3 5 0 3 6 0 3 7 0 3 8 0 3 9 0 4 0 0 410 4 2 0 4 3 0 4 4 0 t, sec

- 4 0

(a) Conventional aerodynamic controls, acceleration command with damping reaction controls, planned CXe = 20°, planned aZ = 4g, ventral off, speed brakes extended 20°, h m a = 226,400 ft.

Figure 13.- Representative entries from high altitude with various types of aerodynamic and reaction controls.

45

h, ft

400

10

ayr g

-1 0

2:[ 2; -20

(b) Adaptive r a t e command controls , planned CXe = 20°, planned aZ = 4g, ven t ra l off , speed brakes extended 20°, haax = 223,700 f t .

Figure 1.3.- Continued.

46

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

1200r soot h, ft - q,

Ib/sq f t

400 t

1 I

I 1 1 I 1 I I 1 I I I I I -1

1, sec

(e) Conventional aerodynamic controls, acceleration command reaction controls, planned Me = 18", planned a2 = 5g, speed brakes extended 35", hmx = 247,000 ft.


47

1200

4 0

*a! deg

- 4 0

Ib/sq f t 400

I . ~

- -.- _ - -

I I I I I I I 1 - L ~

0

2 0

8, deg o

(d) Adaptive rate command controls , a, cp, \I' hold, planned a, = 20°, planned az = 5.5g , hmax = 246,700 f t .


48

I //' 0 L 0 I 1 I -J--- I I I I I I

401- 81 I

sal deg

-40

Q ,A,___,~--'\~)/------~ -----/-,

a,, 9 4 ' .i 0 0 I I I

o---------u----

I I I 1 I I I I I I I

(e) Adaptive rate command controls, planned CXe = 23' , planned az = 5 g , speed brakes extended 20°, hmax = 285,000 ft.


49

IIIIIIII I1 I I I I I

&, d eg

400 0 t

I ,,-- hn- A - -~~ .~ - - _ _ -- -

/----- \ '. ,' q - 2 4 0 -

'. '. / /

/ /

/

/ 80 - /

/

I I I --L -+---< 1 1 I 1 1 - 0

81 I

1 1 I

-1 1 I I I 1 1 I I 1 I 1 I 1

- 4 I-+ -- -- -

/'Q /-

- 2 -

0

I I 1 I 1 1 I I 1 I 1 2

40 I 1

t, sec

(f) Adaptive rate command controls, a, 9, Jr hold, planned Me = 23" , planned aZ = 5g, speed brakes extended 35", hmax = 314,750 ft.


- 4 0 1 I I I I 0

-40

- 2 o L

(g) Adaptive r a t e command controls , planned CXe = 23" , planned aZ = 5g, speed brakes extended 2 O 0 , h" = 347,800 f t .


1l111ll11111l11l I I I I1 I1 I I

400

- ql

Ib/sq ft

a, deg :"i 0

I I I 1 1 I I I I 1 -1 I

/ 1 I I I I I 1 I -I 0

(h) Adaptive rate command controls, planned CXe = 26O, planned a2 = 5.2g, speed brakes extended 20°, hmax = 354,200 f t .


8 0

7 0

60

5 0

Average ae, deg

15

10

I 1 I 3 4 5 6

azmaxl

2 8 0 0

2400

2 0 0 0

1600 - qrnaxr Ib/sq ft

1200

8 0 0

4 0 0

O

(a) hmm = 250,000 f t , V a t hmm = 4,450 ft /sec, airplane weight = 15,000 lb, speed brakes closed.

Figure 14.- Entry recovery requirements for the X-15 .

53

Average ae , deg

15

0 Flight (fig. 13(c))

10

I

- qmaxf Ib/sq f t

1600

1200

800

-1 5

(fig. 13(c))

- 2 0

0 Flight

2 5

400 t "

3

(b) hmax = 250,000 f t , V a t hmax = 4,450 ft /sec, airplane weight = l 5 , O O O lb, speed brakes extended 35'.


7

- qrnaxi

Ib/sq ft

7 0

60

I I 5 0 ' I

2 8 0 0

2400

2000

1600

1200

8 0 0

4 0 0

Average a,, deg

I I 4 5 6

azrnax' 7

( c ) hmax 7 350,000 f t , V a t hmax = 4,260 f t / s ec , a i rp lane weight = 15,000 l b , speed brakes extended 20°.


55

6

5

4

2

1

0

0

0

0

0

0

0

0

Total time above reference-acceleration level , sec

Figure 13.- Sumnary of normal acceleration experienced during entry from high altitude.

I I I I I I 240 280 3 2 0 360 4

I I I 40 80 120 160 200

t, sec 0

6

5

4

M 3

2

I

I

(a) Mach number.

Figure 16.- Summary of entry flight pro f i l e s .

2 4 0 0

2000

1600

- q, 120c

Ib/sq ft

80C

~

0 4 0 8 0 120 160 2 0 0 2 4 0 2 8 0 3 2 0 3 6 0 4 0 0

t, sec

(b) Dynamic pressure.


c'

4 0

32

24

16

8

0

-

- 8 I I I I I I I I 0 4 0 8 0 120 160 2 00 2 4 0 2 8 0 3 2 0 3 6 0 400

t. sec

( e ) Angle of a t t ack .

Figure 16 .- Concluded.

Satisfactory

Unsatisfactory

Satisfactory

Unsatisfactory

Unacceptable

Unacceptable

- 8 -

-

- I I ~~- I J I

c Control system

0 SAS

0 Adaptive

0 Hold

- I I I I I

0 2 4 6 8 10 12

Average a z , g

Satisfactory

Unsatisfactory

Unacceptable

0 0 0

800 1000 1200 1400 1600 1800 2 000

q,,,, Ib/sq ft -

(a) Pitch mode.

Figure 17.- Summary of the p i lo t ra t ing of the X - 1 5 entry control task f o r the three aerodynamic control systems.

60

Satisfactory Control system - 0

0 SAS Unsatisfactory 0 Adaptive

0 Hold -

Unacceptable

- I I I I I

Unsatisfactory

Unacceptable

0 2 4 6 8 10 12

Average a,, g

-

-

- I I I I I

Satisfactory

Unsatisfactory

Unacceptable

8

0 0 -

-

- I I I I I

- 12 16 2 0 24 28 32

Average ae, deg

800 1000 1200 1400 1600 1800 2000

Ib/sq ft - qmaxi

(b) R o l l mode.


61

Satisfactory

Unsatisfactory

Unacceptable

0 2 4 6 8 10 12

Average a = , g

*

Control system - 0 0 SAS

0 Adaptive

0 Hold -

- I I I I I

Satisfactory

Unsatisfactory

Unacceptable

n O 6 0 -

0

-

- I I I I I

( c ) Yaw mode.


Satisfactory

Unsatisfactory

Unacceptable

62

U 0 w 0 - 0 -

- I I I I I

h, ft

32(

28(

24(

20(

160

12a

80

40

Figure 18.- Range required for recovery from an altitude of 334,200 feet.

5

4

3

L/D

2

1

0

Clean configuration

L - - 1 1 1 1 2 3 4 5

M

Figure 19.- Lift -drag r a t i o of t he X-15 with and wTthout speed brakes extended. Power off; a = 10'.

64

3 60x1 0

320

280

2 4 0

2 00

h, f t

160

120

80

40

C

Flight, speed brakes extended 2 0 ° \

- \

L/D,,, clean

configuration

/' i 6 0 ° 6 0 ° bank, Brakes

bank brakes ( 2 0 O ) ( 2 0 " )

40 8 0 120 160 2 0 0 2A

Range, n. mi.

0

Figure 20.- Comparison of f l i g h t range capabili ty with simulator- predicted capabili ty. Vo = 4,310 f t / sec .

Range capo bil i ty,

n. mi.

66

280

240

200

160

120

80

40

0 I I 1 I I I

40 8 0 120 160 200 240

Range to Edwards, n. mi.

2 8 0

Figure 21.- Comparison of actual flight ranging experience with minimum required.

NASA-Langley, 1966 H-384

“The aeronautical and space activities of the United States shall be conditcted so as to contribute . . . to the expansion of hziman Knowl- edge of phenomena in the atmosphere and space. The Administration shall provide for the widest practicable and appropriate dissemination o f information concerning its activities and the results thereof .”

-NATIONAL AERONAUTICS AND SPACE ACT OF 1958

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TECHNICAL REPORTS: important, complete, and a lasting contribution to existing knowledge.

TECHNICAL NOTES: of importance as a contribution to existing knowledge.

TECHNICAL MEMORANDUMS: Information receiving limited distribution because of preliminary data, security classification, or other reasons.

CONTRACTOR REPORTS: Technical information generated in con- nection with a NASA contract or grant and released under NASA auspices.

TECHNICAL TRANSLATIONS: Information published in a foreign language considered to merit NASA distribution in English.

TECHNICAL REPRINTS: Information derived from NASA activities and initially published in the form of journal articles.

SPECIAL PUBLICATIONS Information derived from or of value to NASA activities but not necessarily reporting the results .of individual NASA-programmed scientific efforts. Publications include conference proceedings, monographs, data compilations, handbooks, sourcebooks, and special bibliographies.

Scientific and technical information considered

Information less broad in scope but nevertheless

Details on the availability of these publications may be obtained from:

SCIENTIFIC AND TECHNICAL INFORMATION DIVISION

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summary of high-altitude and entry flight control experience with the

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