National Aeronautics and Space Administration

Langley Research Center • Hampton, Virginia 23681-0001

NASA Technical Paper 3525

Investigation of Piloting Aids for Manual

Control of Hypersonic Maneuvers

David L. Raney, Michael R. Phillips, and Lee H. Person, Jr.

Langley Research Center • Hampton, Virginia

October 1995
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Introduction

Many recently proposed hypersonic vehicle concepts

present significant problems in guidance, flight control,

and flying qualities, which must be addressed if such

designs are to be viable. These air-breathing vehicle

designs traverse a broader range of flight regimes than

aircraft which have flown in the past and must emphasize

performance during all phases of flight to achieve their

mission. Stringent constraints on angle of attack, sideslip, and dynamic pressure must be observed because

hypersonic propulsion system performance will be

strongly dependent on flight condition and may suffer

dramatically if excessive variation is experienced. (See

refs. 1 and 2.) Yet, such vehicles will be required to perform at least several basic maneuvers at high-speed flight

conditions. The anticipated sensitivity of hypersonic propulsion systems makes precise regulation of flight condition imperative, while the high degree of aero-propulsive

interaction makes such tight control difficult to achieve.

These factors, together with nonintuitive effects associated with maneuvering in the hypersonic regime, raise

the question of whether manual control of such a vehicle

should even be considered. Therefore, the extent of manual control which is desirable for a vehicle maneuvering

in this regime must be determined, and the form of aids

that should be supplied to the pilot to make such control

feasible must be identified.

Previous work by Berry (ref. 3) involved a piloted

simulation of a hypersonic configuration in which

numerous handling qualities issues relevant to maneuvering in high-speed flight were examined. Because of propulsion constraints and a number of high-speed flight

effects, pilots encountered considerable difficulty controlling the vehicle even through relatively benign

maneuvers. Subsequent work elaborated on aspects of

the hypersonic flight regime that were responsible for the

observed difficulties and targeted specific areas in need

of additional research. (See refs. 4–6.) These investigations noted that the reduced-g environment of the hypersonic flight regime requires the vehicle to come to bank

angles greater than 60° to execute a level 2g turn. This

effect has been referred to as “centrifugal relief.” Also,

regulation of dynamic pressure during the maneuvers

was particularly difficult because small errors in flight

path angle produced large dynamic pressure excursions

due to density variation with altitude. The dynamic pressure excursions were difficult to arrest because of the

path-attitude decoupling that occurs at hypersonic

speeds. This phenomenon tends to make control of flight

path very difficult if a pitch-rate-command–attitude-hold

response type is used. (See ref. 6.) McRuer and Myers

(ref. 7) further examined the manner in which physical

characteristics of hypersonic air breathers and the associated flight regime would impact flying qualities specifications for this class of vehicle.

Concurrent research by Lallman and Raney (refs. 8–

12) focused on regulation of trajectory parameters during

hypersonic maneuvers at cruise flight conditions, but

maneuvers were restricted to the vertical plane. A

follow-on investigation (ref. 13) addressed the problem

of coordinated control of maneuvers in both the vertical

and lateral planes at hypersonic cruise flight conditions.

An approach was presented that provided maneuver

Abstract

An investigation of piloting aids designed to provide precise maneuver control

for an air-breathing hypersonic vehicle is described. Stringent constraints and

nonintuitive high-speed flight effects associated with maneuvering in the hypersonic

regime raise the question of whether manual control of such a vehicle should even be

considered. The objectives of this research were to determine the extent of manual

control that is desirable for a vehicle maneuvering in this regime and to identify the

form of aids that must be supplied to the pilot to make such control feasible. A piloted

real-time motion-based simulation of a hypersonic vehicle concept was used for this

study, and the investigation focused on a single representative cruise turn maneuver.

Piloting aids, which consisted of an autothrottle, throttle director, autopilot, flight

director, and two head-up display configurations, were developed and evaluated. Two

longitudinal control response types consisting of a rate-command–attitude-hold

system and a load-factor-rate–load-factor-hold system were also compared. The

complete set of piloting aids, which consisted of the autothrottle, throttle director, and

flight director, improved the average Cooper-Harper flying qualities ratings from 8

to 2.6, even though identical inner-loop stability and control augmentation was

provided in all cases. The flight director was determined to be the most critical of

these aids, and the cruise turn maneuver was unachievable to adequate performance

specifications in the absence of this flight director.
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coordination through the resolution of altitude and crossrange errors into a normal load factor and bank angle

command combination. The research of reference 13

yielded an automatic control design for executing coordinated maneuvers in hypersonic flight while regulating

key parameters such as angle of attack, sideslip, and

dynamic pressure.

This report describes an investigation in which the

control laws and guidance concepts presented in reference 13 were implemented in a real-time simulation of an

air-breathing hypersonic vehicle with the objective of

development and evaluation of a set of piloting aids that

could provide desired maneuver performance with an

acceptable pilot workload. The real-time experiment

consisted of two parts. The first part of the experiment

was intended to identify the pilots’ preferences from a set

of two response types and two head-up display configurations designed to permit satisfactory manual control of

the hypersonic vehicle. The second part of the experiment examined the degradation in task performance and

pilot satisfaction as various pilot aids were removed and

compared the resulting performance with maneuver performance in the fully automatic mode. The experiment

provided information regarding the level of assistance

that must be supplied to the pilot to achieve desired

maneuver constraints and the form that the assistance

should take.

The report begins with a description of the aircraft

model used in the real-time simulation. The physical

configuration of the simulator itself, in terms of displays

and control inceptors, is briefly presented. Next, the

report describes the real-time experiment design,

describes the evaluation criteria, and discusses the experiment results. A concluding section summarizes the

observations and presents several implications for vehicles maneuvering in the hypersonic regime.

Symbols

g acceleration due to gravity, ft/sec2

h altitude, ft

n normal load factor, g units

dynamic pressure, lb/ft2

R altitude measured to center of Earth, ft

t

θ2 flight path-attitude lag time constant

V total velocity relative to air mass, ft/sec

α angle of attack, deg

γ flight path angle, deg

∆ perturbation from trim conditions

ρ density, slug/ft3

φ bank angle, deg

Abbreviations:

APAS aerodynamic preliminary analysis system

CDU cockpit display unit

CGI computer-generated image

CHR Cooper-Harper rating

EADI electronic attitude display indicator

HSI horizontal situation indicator

HUD head-up display

LFRC load factor rate command

PAL pilot assistance level

POST Program to Optimize Simulated Trajectories

RBFD resolver-based flight director

RCAH rate-command–attitude-hold

rms root-mean-square

SPI surface position indicator

VMS Visual/Motion Simulator

Aircraft Model

The vehicle concept used in this study is an airbreathing single-stage-to-orbit configuration designed to

operate from a conventional runway. Planform and profile views of this hypersonic winged cone configuration

are presented in figure 1. The fuselage has a conical forebody half angle of 5° with wrap-around engine nacelles

and a 75° swept delta wing. The configuration includes

elevons and a rudder for longitudinal and lateraldirectional control. The overall vehicle length is 200 ft

with a wing span of 60 ft; the weight of the fully fueled

vehicle is approximately 300 000 lb with a fuel load of

110 000 lb. The pilot is located 101 ft ahead of the vehicle center of gravity at takeoff. A summary of geometric

characteristics for this vehicle is presented in table 1.

A full-envelope database has been generated for the

configuration as a result of extensive wind tunnel testing

combined with analytical investigations using the aerodynamic preliminary analysis system (APAS). (See

ref. 14.) The development of this inertial, propulsive, and

aerodynamic database is detailed in references 15 and 16.

The data provided in reference 15 were used to represent

the aerodynamic characteristics of this configuration for

the real-time simulation.

The variation of inertial characteristics and center of

gravity with total vehicle weight is shown in figure 2. All

maneuvers presented in this report were initiated with a

vehicle weight of 250 000 lb. Although mass reduction

due to fuel burn was modeled, inertial characteristics and

q
3

center-of-gravity location did not vary significantly during the real-time maneuver simulations.

The engine model used in this investigation does not

include propulsion sensitivities to angle-of-attack variation. For this particular configuration, the wrap-around

engine nacelles may tend to reduce the effect of propulsion sensitivity to angle of attack, but more typical configurations often include an underslung engine nacelle

located on the lower surface of the fuselage. Such

configurations would likely exhibit a greater propulsion

sensitivity to angle-of-attack variation and thereby complicate longitudinal control of the vehicle. (See ref. 17.)

Details of the engine model are presented in appendix A.

The simulation model used in this study is a rigidbody approximation, although recent analytical investigations using other hypersonic configurations have indicated that the aeroelastic modes have significant

potential impact on all aspects of control for this type of

vehicle. (See refs. 18 and 19.) The impact of aeroelasticity on the manual control of such configurations is a

topic worthy of real-time research, but such a study

would require a motion-based simulator with a bandwidth that exceeds the capabilities of the facility used in

this investigation.

Simulation Facility

A real-time simulation of the hypersonic aircraft

model was implemented on the NASA Langley Visual/

Motion Simulator (VMS). The Langley VMS is a sixdegree-of-freedom hexapod with performance characteristics described in table 2. During this investigation, the

approximate time delay for the motion cues was 82 msec

(+30 msec) and for the visual cues was 170 msec

(+35 msec). Detailed specifications of the Langley

VMS facility and motion algorithms are provided in references 20 and 21. The spherical rotational Earth equations of motion that were used to drive the simulator

visual displays and motion cues are drawn primarily

from reference 22. The general interior layout of the

cockpit is shown in figure 3. The test pilot occupied the

left station and the research engineer occupied the right

station for all of the maneuvers described in this report.

Descriptions of individual control and display features

shown in figure 3 follow.

HUD and Out-the-Window Scene

The out-the-window scene used for the visual component of this simulation was a computer-generated

image (CGI) produced by an Evans & Sutherland CT-6

system. Because the maneuvers in this investigation took

place at an altitude of approximately 85 000 ft, the outthe-window visual scene produced by the CGI system

simply consisted of dark blue sky showing a cloud layer

to provide an indication of roll attitude with respect to

the horizon. The image was presented on three cockpit

windows (two forward and one left side) with an arrangement of three video monitors and a mirror beam splitter.

Further details regarding the physical arrangement of the

components of this display system are available in

reference 23.

The primary cockpit flight instrument used in this

investigation consisted of a head-up display (HUD). The

HUD image was generated by a Terabit Eagle 100 symbology generator and then superimposed on the out-thewindow scene generated by the Evans & Sutherland CGI

system. This combined image was then presented to the

pilot on the forward cockpit window video monitor.

Although the HUD contained all information necessary

to perform the hypersonic maneuver tasks, pilots indicated that inclusion of the out-the-window scene provided reassurance of the vehicle attitude with respect to

the horizon.

The baseline HUD, which was used in this experiment, is shown in figure 4 and contains information relevant to the flight condition and vehicle attitude. A

conventional pitch ladder with attitude indicator is provided at the center of the display. An artificial horizon

line and flight path marker are also included as shown in

figure 4. To the right of the horizon line is a digital altitude indicator and to the left of the horizon line are digital indicators for equivalent airspeed, angle of attack,

Mach number, dynamic pressure, and normal load factor.

A bank angle index is provided in the form of radial tick

marks distributed in an arc around the bottom of the display in increments of 30°. As shown in figure 4, minor

tick marks occur in increments of 10° between bank

angle indices of ±30°. A triangular symbol travels around

the perimeter of the bank angle index to indicate the

vehicle roll attitude. At the top of the display is a heading

indicator and index with tick marks spaced at intervals of

2° and labeled at intervals of 10°. A throttle level indicator depicts the current throttle setting on a vertical scale

shown in the upper left corner of figure 4.

Two radically different forms of flight director symbology were used with this basic HUD layout. As part of

this experiment, the pilot’s ability to perform the hypersonic maneuver task with different types of director symbology was compared. Specific components of these two

flight director formats and the algorithm used to drive

them will be discussed later in this report.

Head-Down Displays

Figure 3 shows the location of the electronic attitude

display indicator (EADI) on the cockpit instrument

panel. The EADI contains much of the same information

presented on the basic HUD previously described but
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without the presence of the out-the-window scene. A diagram of the EADI is shown in figure 5. The pilots who

participated in this experiment indicated that they rarely,

if ever, referred to the EADI during the maneuvers, and

then, only to obtain gross confirmation of the information being displayed on the HUD.

Figure 3 also shows the location of the horizontal situation indicator (HSI) and surface position indicator

(SPI). The HSI is shown in figure 6. Only the compass

and heading readouts of this display were driven during

this experiment, and pilots indicated that this instrument

was not needed during the execution of the up-and-away

maneuvers as long as they could rely on the heading

index shown on the HUD.

A diagram of the SPI, which is shown in figure 7,

depicts the magnitude and sense of the various control

surface deflections with a color code for rate or position

saturation. The control surfaces are colored green during

nominal operation. If an actuator experiences a rate limit,

the corresponding control surface is colored yellow; if a

position limit is experienced, the surface turns red. Readouts of fuel flow rate and total fuel consumption are

included on this instrument as well as a fill line, which

shades the fuselage to provide a general indication of the

amount of fuel remaining. This instrument was used primarily by the research engineer during control system

checkout and diagnostic activities. Pilots did not use this

display during execution of the maneuvers.

Cockpit Controls

The cockpit controls available to the pilot consisted

of a left-side stick controller and a single throttle lever.

The general location of the side stick is shown in

figure 3. Force gradient and breakout settings of the

McFadden hydraulic control loader used for this side

stick are shown in table 3. An adjustable armrest was

provided adjacent to the controller, and pilots indicated

that no discomfort or awkwardness was experienced with

the use of this type of control inceptor.

The throttle lever controlling total fuel flow rate was

located to the pilot’s right as shown in figure 3. The pilot

of such a vehicle would not likely exercise independent

control over fuel flow rate to individual combustor modules, so a single throttle lever was provided to the pilot

for this study. An aural cue was provided in the form of

engine noise that was scaled with fuel flow rate. (The

noise was actually an afterburner sound effect, so turbine

spooling sounds were intentionally absent.) The throttle

lever had a total throw of 52.5°, and the corresponding

fuel flow rate varied from 10 lb/sec in the aft position to

200 lb/sec in the full-forward position. Appendix A illustrates the manner in which the commanded fuel flow rate

was used in the engine model.

The level of throttleability that would be possible for

a scramjet propulsion system operating at hypersonic

speeds is uncertain. Uninterrupted operation of the

propulsion system would be of paramount importance

during a hypersonic maneuver, which would place constraints on allowable magnitudes and rates of change of

fuel flow settings. Such constraints will depend on configuration and flight condition. However, the severity of

these constraints is highly uncertain, and information is

not currently available on the throttleability of potential

hypersonic propulsion systems. In the interest of ascertaining what the pilot would prefer to do in the absence

of constraints, no limits were placed on allowable throttle

lever angle variation when the throttle was under manual

control.

Additionally, a cockpit display unit (CDU) was used

by the research engineer to configure the simulator. The

location of this device is shown in figure 3. The CDU

was programmed to permit the researcher to select command response type, flight director type, and pilot assistance level (PAL) from the cockpit prior to each

maneuver. (Pilot assistance level refers to the number

and type of aids supplied to the pilot during the maneuver and will be described in greater detail later in this

report.) Specific target flight conditions for the maneuver

were also input to the CDU, and initiation of the task was

manually triggered from the CDU. The CDU program

included maneuver task completion criteria (table 4) that

could provide a visual cue upon successful completion of

a maneuver. The exact values from table 4 were selected

somewhat arbitrarily, but they still provided a consistent

definition of task completion for use when comparing

time histories of a given maneuver under differing control configurations.

Preliminary Simulation Sessions

At the outset of this investigation, a series of realtime simulation sessions were conducted in which two

Langley test pilots performed several of the hypersonic

maneuver tasks described in reference 3. These consisted

of constant dynamic pressure ascents and descents,

decelerating turns, and cruise turns. The control system

was a rate-command–attitude-hold response type for

both the pitch and roll axes. (The design of this control

law is described in appendix B.) No flight director or

autothrottle was provided, but the baseline HUD shown

in figure 4 was presented. When attempting to perform

the maneuvers, the pilots experienced numerous difficulties associated with the peculiarities of the hypersonic

flight regime that were revealed during earlier hypersonic simulation work conducted by Berry. (See refs. 3

and 4.) The intent of these preliminary simulation sessions was to reproduce these effects so that potential

solutions to the problems could be developed. The
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hypersonic cruise turn maneuver was selected for this

experiment because it required coordination of lateral

and longitudinal inputs as well as considerable throttle

activity. To the test pilots, this task represented a fundamental maneuver with a high degree of difficulty.

The hypersonic cruise turn maneuver consisted of a

heading change of 30° at constant altitude and dynamic

pressure and was executed at a load factor of 2g. The

maneuver was initiated from a level trimmed flight condition at Mach 7.86 at an altitude of 85040 ft and

dynamic pressure of 2000 lb/ft2. The normal load factor

required to maintain straight and level flight at this condition was approximately 0.84g because of the centrifugal relief effect. Centrifugal relief refers to a reduction in

the normal load factor required for level flight that is

experienced as a vehicle approaches orbital velocity.

(See refs. 4 and 5.) The hypersonic cruise turn task definition was adapted from reference 3 and is summarized

as follows:

1. Hold Mach number, altitude, and dynamic

pressure

2. Bank to initiate maneuver

3. Establish steady 2g turn

4. Attain required heading change

5. Roll-out on desired heading

When asked to perform this maneuver with no flight

director or autothrottle, the pilots indicated that the task

was not reliably achievable. Note that the pilots felt that

they had good control over the aircraft attitude; the ratecommand–attitude-hold control law was doing its job.

Figure 8 shows typical time histories of the cruise turn

executed by two different pilots using the ratecommand–attitude-hold system without autothrottles or

flight director guidance. The target heading for this

maneuver was 120°. From the heading time histories

shown in figure 8, the pilots clearly did not have precise

control over this parameter. The bank angle traces also

show that there was a considerable uncertainty associated

with the timing of the roll-out needed to arrest turn rate at

the target heading.

The target altitude for the cruise turn was 85 000 ft

and dynamic pressure was 2000 lb/ft2. Figure 8 shows

that significant excursions also occurred in these parameters during the maneuver. The pilots apparently were not

able to control flight path sufficiently to perform the

desired maneuver. Both pilots noted that arresting flight

path deviations during the cruise turn was particularly

difficult. The primary difficulty arises from the decoupling of path and attitude that occurs for this vehicle at

the chosen flight condition. This result was not unexpected because it has been repeatedly alluded to in earlier research (refs. 3, 5, and 6). In fact, the development

by Chalk in reference 6 regarding rate of change of flight

path for a hypersonic vehicle equipped with a pitch ratecommand–attitude-hold system predicted that extremely

low-bandwidth control of flight path rate would result.

The reason for this is illustrated by equations (1) and (2).

(1)

(2)

where ∆n is the normal load factor perturbation from

trim.

The parameter 1/t

θ2

 is referred to as the flight pathattitude lag and can be interpreted as the frequency above

which ∆θ = ∆α and below which ∆θ = ∆γ. (See ref. 6.)

Note in equation (1) that, for a given n/α, 1/t

θ2

 tends to

zero as V increases. Equation (2) further illustrates that,

for a given change in normal load factor, the resulting

rate of change in flight path angle is inversely proportional to velocity. Therefore, pitch attitude changes in

hypersonic flight take an excessively long time to produce flight path changes. (The value of 1/t

θ2 for the subject vehicle at the Mach 7.86 flight condition is

approximately 0.084 sec−1.) This effect is referred to as

“path-attitude decoupling.” The plots of pitch attitude

and flight path angle time histories shown in figure 9

illustrate the impact of this decoupling. This figure

shows that flight path angle was relatively unaffected by

the pilots’ higher frequency pitch inputs, while the flight

path response to the pilots’ lower frequency pitch inputs

was virtually 90° out of phase. The resulting angle-ofattack perturbations shown in figure 8 are clearly not

desirable from the standpoint of minimizing disturbances

to the propulsion inlet conditions.

Equation (3) further illustrates the importance of

flight path in regulating dynamic pressure. At hypersonic

speeds, the V3 multiplier of density gradient becomes the

dominant factor in this equation. As flight path errors

developed during the turn, the pilots noted a dynamic

pressure error, which they then attempted to null with the

use of throttle inputs. This significantly increased pilot

workload during the maneuver. The time histories in

figure 8 show that pilot B increased fuel flow rate in an

unsuccessful attempt to compensate for dynamic pressure errors resulting from a positive altitude excursion.

(3)

The next section of this report describes several

piloting aids that were developed to address the

1

t

θ2

------ g

V

--- n

α = ---

γ ̇ g∆n

V = ---------

q ̇ ρV V ̇ V3

2

------ ∂ρ

∂h = + ----- γ
6

difficulties that the pilots experienced during the preliminary simulation sessions.

Description of Piloting Aids

The difficulties encountered during manual control

of the cruise turn maneuver precipitated an informal

development phase in which the research engineer

worked together with the project pilots to develop a

series of aids in the form of guidance laws, flight directors, autothrottle, and throttle director that would help the

pilots cope with the hypersonic flight environment. This

part of the research involved discussions with the pilots,

development and implementation of solutions, and precursory evaluation of these solutions leading to further

refinement of the concepts. These piloting aids were then

systematically evaluated in a formal real-time experiment. The following paragraphs describe the aids that

were developed to assist the pilots in conducting the

hypersonic maneuvers. The formal experiment that was

designed to evaluate and quantify the benefits derived

from use of these aids will be described later in this

report.

Autopilot and Flight Director Guidance

Trajectory control was recognized to be the primary

difficulty that the pilots would encounter during the

cruise turn. Chalk noted in reference 6 that direct piloted

control of flight path will not be practical because of the

slow response rate; a better approach may be for the pilot

to exert direct control over angle of attack or load factor

and to modulate his inputs in response to flight director

displays for trajectory control. Later work by Raney and

Lallman (ref. 13) acknowledged this and addressed the

problem by using a guidance approach that resolved vertical and lateral trajectory deviations into a lift vector

command that would null the errors.

The primary element of the guidance law developed

in reference 13 was a simple coordinate transformation

referred to as a “resolver” that converted vertical and lateral flight path corrections into a desired bank angle and

normal load factor subject to vehicle constraints. The

guidance law was applied to an example configuration

identical to the configuration used in this study, and its

performance was demonstrated in an unpiloted digital

simulation. The algorithm could be used either to drive

inner-loop flight controls directly as an autopilot or to

drive a cockpit flight director that could guide the pilot

through a hypersonic maneuver. The guidance law was

implemented in the real-time simulation for this study so

that it could operate in either of these two modes.

Detailed development of this guidance strategy is presented in reference 13.

The advantage of this guidance concept for driving a

flight director, rather than simply presenting a flight path

error to the pilot, is that flight path dynamics are effectively removed from the immediate control task that the

pilot is attempting to perform. This is achieved because

the guidance law specifies flight director cues in the form

of normal load factor and bank angle, which are two

parameters over which the pilot has relatively high-bandwidth control even at hypersonic flight conditions. The

effect of path-attitude decoupling is thereby eliminated

from the pilot’s flight director tracking task, and the

problems of a conventional flight path director are

avoided. This is precisely the approach envisioned by

Chalk in reference 6.

Flight Director Symbology

Two forms of flight director symbology were compared for use in executing the hypersonic maneuvers.

The two flight directors differed only in the graphical

presentation of the load factor and bank angle cues to the

pilot. The first director symbology is referred to as the

“conventional flight director” in this report, and the second is referred to as the “resolver-based flight director”

(RBFD). The same algorithm was used to compute the

desired load factor and bank angle in both cases.

The symbology for the conventional flight director

used in this investigation consists of a director diamond

superimposed on the head-up display from figure 1 as

shown in figure 10. The vertical offset of the director

symbol from the flight path marker is driven by the normal load factor error, which is the difference between the

actual load factor normal to the velocity vector and the

desired load factor computed according to the algorithm

described in reference 13. Likewise, the lateral offset of

the flight director symbol from the flight path marker is

driven by the bank angle error, which is the difference

between the actual bank angle of the vehicle and the

desired bank angle generated by the guidance law. The

director diamond in figure 10 flashes from hollow to

solid for 2 sec prior to the initiation of the roll-out during

the cruise turn maneuver.

Because the flight path angle was no longer the

parameter being directly pursued by the pilot, revision of

the visual cue used to depict the control task was also

appropriate. The resolver-based flight director (RBFD)

presents the pilot with a graphical depiction of the magnitude and direction of the vehicle actual lift vector

instead of the desired lift vector specified by the guidance law. The display is simply a pictorial representation

of the resolver element of the guidance algorithm

described in reference 13.

Primary elements of the RBFD symbology consist of

the load factor director circle, bank angle director, and
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the lift vector arrow shown in figure 11. The load factor

circle is centered on the flight path marker symbol. The

radius of the load factor circle is proportional to the magnitude of the desired load factor computed by the

resolver element of the guidance algorithm. The load factor circle contracts or expands to indicate the variation of

the desired load factor. The bank angle director symbol

moves around the perimeter of the load factor circle to

indicate the desired bank angle computed by the resolver.

A faint dashed line runs from the center of the flight path

marker to the bank angle director symbol. A lift vector

arrow is shown in figure 11 pointing vertically upward

from the origin of the flight path marker symbol. The

length of the lift vector arrow is proportional to the actual

load factor acting normal to the vehicle flight path based

on accelerometer measurements. The lift vector arrow

always points straight up from the flight path marker.

To execute a desired maneuver with the RBFD, the

pilot manipulates the control stick to place the tip of the

lift vector arrow at the location on the perimeter of the

load factor circle indicated by the bank angle director. In

this way the pilot achieves the load factor and bank angle

that will cause the flight path to move in the desired fashion. As the desired load factor and bank angle vary, the

pilot tracks the RBFD symbology by attempting to maintain contact between the tip of the lift vector arrow and

the bank angle director on the perimeter of the load factor

circle. The dashed line between the bank angle director

and the center of the flight path marker permits the pilot

to ascertain whether he has precisely matched the desired

bank angle. The dashed line will be completely obscured

by the lift vector arrow when the bank angle error is

nulled.

After the two test pilots performed several maneuvers with the RBFD, they suggested a refinement to the

symbology, which resulted in a fixed-radius version of

the RBFD shown in figure 11. The load factor circle was

fixed at a constant radius, and the lift vector arrow shortened or lengthened to reflect the error between the

desired load factor and the commanded load factor rather

than having the load factor circle contract or expand

according to desired load factor. The test pilots indicated

that too much appeared to be happening at the same time

when the load factor director circle and the load factor

arrow were both changing size in the original RBFD.

The nominal bank angle indicator shown in figure 11

is another refinement based on pilot suggestions and is

used only during maneuvers for which a prescribed nominal bank angle can be defined; the level 2g turn is such a

maneuver. In this case, the nominal bank angle indicator

appears at the nominal bank angle for the turn on the

inner perimeter of the load factor circle at the instant the

maneuver is initiated. The bank angle director then

moves around the outer perimeter of the load factor circle

at the desired roll rate until it is aligned with the nominal

bank angle indicator. The pilot tracks the bank angle

director during the roll-in. The nominal bank angle for a

level 2g turn at hypersonic speeds is calculated from

equation (4).

(4)

During the turn, the bank angle director may call for

minor deviations from the nominal bank angle indicator

to null slight altitude errors. The nominal bank angle

indicator symbol was added to provide the pilots with an

anticipation cue that would warn them to prepare to

arrest their roll rate. The bank angle director triangle in

figure 11 flashes from hollow to solid for 2 sec prior to

the initiation of the roll-out during the cruise turn

maneuver.

The specific symbology of the RBFD is designed to

represent the physical orientation of the vehicle lift vector in relation to the desired load factor and bank angle.

In this way the RBFD relates directly to the fundamental

forces, which act on the aircraft and which the pilot must

manipulate in order to effectively control the trajectory.

Command Response Types

Two command response types were compared for

use in executing the hypersonic maneuver tasks. The first

response type was a pitch rate-command–attitude-hold

(RCAH) system. This is the same system that was used

to perform the maneuvers during the preliminary simulation sessions. Because the guidance laws were designed

to drive the flight director based on desired load factor,

the pilots were then provided with another response type

option that might provide better harmony with such a

director. Accordingly, a second control law was designed

to provide a load-factor-rate-command (LFRC)–loadfactor-hold response type. Note that such a system would

be self-trimming in the sustained 2g turn and should

require less pilot attention than the rate-command–

attitude-hold system. The LFRC system was also well

suited to the incorporation of load factor limits. In both

cases, the lateral control law was a roll-rate-command–

attitude-hold system. A description of the control laws

for the two response types is provided in appendix B.

The set of two flight director symbologies and two

response types yielded four combinations for use in the

piloted evaluations. The combination of the conventional

director symbology with the RCAH response type represented a relatively familiar configuration to the test

pilots, while the combination of the resolver-based flight

φnom cos–1 0.5 1 V2

Rg – ------     =
8

director and the LFRC response type represented a considerable departure from convention.

Autothrottle and Throttle Director

Another piloting aid that was designed for this investigation consisted of an autothrottle that regulates

dynamic pressure during hypersonic maneuvers. Pilots

had indicated a high degree of difficulty associated with

regulating dynamic pressure during the cruise turns in the

preliminary simulation sessions. An autothrottle, which

compensated for density variation with altitude, was formulated to maintain dynamic pressure during ascending

or descending flight. The design of this autothrottle was

developed in reference 13.

A throttle director was also provided on the left

wingtip of the flight path marker to provide a speed control cue in the absence of the autothrottle system. This

element of the display is shown in figures 10 and 11. The

throttle director consists of a shaded bar which appears

above or below the left wingtip of the flight path marker.

The vertical length of the bar is proportional to the fuel

flow error, which is the difference between the actual

fuel flow rate and the desired fuel flow rate to the propulsion system. The desired fuel flow rate is computed with

the same algorithm that drives the autothrottle. The bar

drops below the wingtip to indicate that the actual fuel

flow rate is less than the desired rate and rises above the

wingtip when the actual fuel flow rate is greater than the

desired rate. The pilot is then able to null the fuel flow

error by adjusting the throttle controls accordingly. An

identical throttle director was used with the conventional

flight director symbology and the RBFD symbology.

Pilot Assistance Levels

 To ascertain the form and degree of assistance that

must be supplied to the pilot to achieve satisfactory

maneuver performance, the pilot assistance level (PAL)

concept was developed. The numerical PAL designation

represents the level of assistance provided to the pilot in

the form of directors and automation. A flowchart that

depicts the manner in which the various piloting aids

were integrated with other elements of the real-time simulation is shown in figure 12. Table 5 defines the set of

numerical PAL designations and presents the associated

switch configurations for the flowchart. The autopilot

and autothrottle elements shown in figure 12 are thoroughly described and demonstrated in reference 13. A

description of the inner-loop pitch, roll, and yaw control

elements is presented in appendix B. Figure 12 shows

that the basic inner-loop stability augmentation system

(SAS) was provided with all PAL conditions and that

only outer-loop elements were added or removed as the

pilot assistance level was varied.

Prior to initiation of a run, the pilot assistance level

was selected by entering the appropriate numerical PAL

designation from table 5 into the CDU. This process

selected which piloting aids would be supplied by appropriately configuring the switches shown in the flowchart.

The CDU was then used to select which of the two

HUD’s and two longitudinal control response types

would be used.

After designating the PAL setting, the maneuver task

was input to the CDU by specifying desired changes in

altitude, dynamic pressure, and heading. The cruise turns

that were performed in this study consisted of a heading

change of 30° at constant altitude and dynamic pressure.

The maneuver was then initiated by pressing the “execute” button on the CDU. The guidance laws then drove

the flight director, autothrottle, or autopilot corresponding to the PAL setting. If the PAL setting was 0, then neither the autopilot, flight director, autothrottle, nor throttle

director was driven, and the pilot had full manual control

of the vehicle using the side-arm control stick and throttle levers.

From the comparison of performance metrics for a

given task executed with different PAL configurations,

identification of those tasks which require automation or

flight director cues to achieve the desired performance is

possible. This information is necessary to define the

types of tasks that can and cannot be performed under

degraded control configurations and to determine the

level of automation that should be provided for controlling the vehicle during nominal execution of a hypersonic maneuver.

Experiment Design

The hypersonic cruise turn maneuver was used to

compare the various control designs and piloting aids in a

formal real-time experiment. When comparing the control and flight director features, the pilot was asked to

focus on the following two main issues: Does the feature

improve the ease with which the task is executed? and

Does the feature improve the precision with which the

task is executed? After each task, the pilot was asked to

evaluate the vehicle control systems using the CooperHarper flying qualities rating scale shown in figure 13.

Subjective comments on displays and response types

were also requested. During each run, a videotape of the

front cockpit monitor showing the CGI and HUD was

made, which included an audio recording of the pilot’s

remarks.

The real-time experiment consisted of two segments.

Part 1 of the experiment was designed to identify the

pilot’s preferred command response type and flight

director symbology from a set of four combinations for
9

use in the hypersonic cruise turn maneuver. The designations for these four response type and HUD combinations

are shown in table 6. Part 1 of the experiment was conducted with PAL 3 so that autothrottles were active at all

times. This permitted the pilot to focus on the specific

features of the response type and the flight director symbology. At the conclusion of part 1, the pilot was asked

to select his preferred flight director and response type

combination.

Part 2 of the experiment was intended to assess the

degradation in task performance and Cooper-Harper

flying qualities ratings as the pilot assistance level was

reduced from 4 (fully automated) to 0 (no automation).

This segment of the experiment was conducted with the

pilot’s preferred director symbology and command

response type from part 1. When the maneuver was executed with PAL 4, the pilots simply observed the automatic systems while paying special attention to the

timing and magnitude of throttle inputs. This provided

them with an opportunity to train for the task of conducting the maneuver when the autothrottle and flight director would not be present.

Four Langley test pilots were asked to participate in

the experiment. All four pilots were familiar with a wide

range of configurations including both transport and

fighter aircraft. Each pilot was briefed on the purpose of

the experiment prior to participation. This briefing also

included a discussion of the flying qualities and control

issues associated with maneuvering in hypersonic flight.

Prior to each simulation session, the pilot was supplied

with a specific session itinerary, run log, and comment

pages. These materials described the maneuver tasks and

objectives of the simulation session. A typical session

itinerary is shown in figure 14.

Specific criteria were necessary for judgment of

desired and adequate task performance. Information on

tolerances of hypersonic propulsion systems to angle-ofattack and dynamic pressure perturbations and permissible range of throttleability was not available at the time

of this investigation. Therefore, relatively stringent performance criteria were specified in the interest of providing a conservative assessment of the errors that the

propulsion system and trajectory parameters might tolerate. Figure 14 includes the criteria used to define desired

and adequate task performance for dynamic pressure and

altitude regulation as well as heading capture. Note that

neither of the time histories shown in figure 8 satisfied

even the adequate performance criteria. Desired and adequate tolerances for the tracking of guidance commands

were also specified to let the pilots know how aggressively they should pursue the director symbology.

To assess the merits of the various piloting aids,

maneuver tasks were compared on the basis of two criteria. The first means of comparison was pilot workload

metrics, and the second was task performance metrics.

These metrics are listed as follows:

1. Pilot workload metrics

A. Flying qualities ratings using Cooper-Harper scale

B. Longitudinal stick rms deviation from mean

C. Flight director rms tracking errors

D. Subjective comments

2. Performance metrics

A. Peak absolute error measurements in dynamic

pressure and altitude

B. Dynamic pressure and altitude rms error

C. Fuel consumption during task

D. Acquisition and regulation of target parameters

The pilot workload metrics addressed the question of

whether or not a particular feature being examined

improved the ease with which the pilot was able to perform a given maneuver, while the performance metrics

addressed the question of whether or not the feature

improved the precision with which the pilot could perform the maneuver. The following section of the report

describes the results of the real-time experiment and presents comparisons in terms of these metrics.

Discussion of Part 1 Results

The first part of the experiment was intended to identify the pilots’ preferences for the response type and

HUD combination. Each pilot performed the hypersonic

cruise turn maneuver several times with each of the four

combinations. The authors emphasize that this part of the

experiment was conducted with the use of PAL 3 in all

cases, that is, with the autothrottle activated. Table 7

shows the pilots’ final choices for their preferred

response type and HUD symbology. A clear preference

was demonstrated for the LFRC response type over the

RCAH. Pilots indicated that the workload was much

lower with this response type but that the need to push

the stick forward in the roll-out to unload the vehicle was

disconcerting at first. The harmony of the flight director

with this response type was understandably much greater

because the flight director symbology for either HUD

configuration is driven by a load factor error signal.

Table 7 shows no consistent preference for one flight

director symbology over the other, although each pilot

cited significant reasons for his choice. Appendix C contains a transcript of each pilot’s comments regarding the

basis for the selection of the preferred response type and

HUD combination. The following sections examine the
10

variation in pilot workload metrics and task performance

metrics with the response type and the HUD symbology.

Workload Metrics

The test pilots were asked to assign a Cooper-Harper

flying qualities rating (CHR) to each of the response type

and HUD combinations. These ratings are shown in figure 15. The data show an improvement in pilot satisfaction with the LFRC response type when compared with

the RCAH response type. The pilots generally indicated

that the LFRC control system permitted the flight director signal to be followed with greater ease. Pilots also

commented that the LFRC response type provided

greater ability to carefully control and arrest the g-onset

than the RCAH system during roll-in and roll-out

maneuvers.

Figure 15 shows that pilot C awarded the LFRC–

HUD-1 combination a CHR of 2.5, while awarding the

LFRC–HUD-2 combination a CHR of 4. Pilot C

expressly noted that the source of this rating difference

was a hardware issue and felt the basic symbology elements of HUD 2 provided the pilot with better awareness

than HUD 1. However, as the large circular element of

HUD 2 (the load factor circle shown in fig. 11) traversed

the horizontal scan lines of the cockpit window video

monitor, the curved line appeared to stair-step. This

effect proved particularly distracting to pilot C, which

resulted in lowering the rating of HUD 2. This effect is

an artifact of the hardware configuration of the Langley

VMS, which mixes the HUD with an out-the-window

scene for display on a video monitor. Pilot C indicated

that, although the conceptual depiction of the task was

better with HUD 2 (the RBFD), the higher resolution of

the display symbols provided by the HUD 1 implementation was preferred.

The combination that received the best average CHR

was the LFRC response type with the HUD-1 symbology. Each of the test pilots awarded this combination a

level-1 flying qualities rating (CHR < 3). As table 7 indicated, this combination was selected as the most preferable by two of the test pilots.

Figure 16 shows a comparison of the rms longitudinal stick deflections associated with each of the four

response types and HUD combinations for the cruise turn

task. To produce this graph, the rms deviation from the

mean stick deflection was computed for each of the time

histories in the entire set of real-time runs conducted by

all four test pilots. The average rms stick deflection was

then computed for each response type and HUD combination as shown in figure 16. The maximum and minimum rms deflections are also shown in the figure. The

values represent the general level of stick activity about

the mean stick deflection. Approximately twice as much

stick activity resulted from the RCAH system than from

the LFRC system. However, a few comments regarding

possible bias in this observation should be noted.

First, the basic nature of the cruise turn task, which

requires the pilot to maintain 2g through the turn while

regulating altitude (and thereby dynamic pressure) with

bank angle, makes the LFRC response type ideally suited

to this task. Other tasks not requiring sustained periods of

constant g may have resulted in less pronounced differences in stick activity between the RCAH and LFRC

control systems. Also, the LFRC response type permits

the pilot to carefully control the rate of g-onset, set load

factor, and then get out of the longitudinal control loop

during the steady segment of the maneuver. But the

RCAH system did not include a trim control and, therefore, required the pilot’s constant attention to maintain

2g during the steady segment of the maneuver. Inclusion

of a trim button on the stick for use with the RCAH system might have reduced this effect, although the pilots

commented that they would be hesitant to set such a trim

device at the value required for the 2g turn.

Further insight into these effects is provided in

figure 17, which presents a comparison of time histories

from typical maneuvers conducted by pilot A using each

of the four response type and HUD configurations. The

figure shows that similar levels of altitude and dynamic

pressure errors were experienced for all four configurations, and these errors were well within the tolerances

specified for desired performance. However, figure 17

also shows that normal load factor was relatively quiet

during the steady segment of the maneuver for the runs

conducted with the LFRC response type and that considerably greater activity was experienced with the RCAH

system. Figure 17 also shows the associated angle-ofattack time histories. Although the magnitudes of the

angle-of-attack variation during the maneuvers are the

same, the time histories show that the runs conducted

with the RCAH system were much noisier. The pilots

commented that the RCAH response type required continual inputs on their part and that the precision of the

maneuver was not as good as with the LFRC system.

Anticipated propulsion system sensitivity to angle of

attack makes minimization of perturbations during the

maneuver desirable; the benefit of the LFRC response

type, which allows the pilot to extract himself from the

longitudinal task during the steady segment of the

maneuver, is clear.

Figure 17 also shows that about the same level of

bank angle activity was experienced during the maneuver

for all four configurations. The small bank angle perturbations are called for by the guidance algorithm to regulate altitude and dynamic pressure during the maneuver.

A comparison of load factor and bank angle time
11

histories in figure 17 clearly indicates the necessity to

roll and pull simultaneously to track flight director guidance during the roll-in and roll-out segments of this

maneuver.

A comparison of the heading, altitude, and dynamic

pressure time histories in figure 17 with those in figure 8

clearly shows that use of the flight director and guidance

algorithm developed after the preliminary simulation sessions provided the pilot with significantly greater control

over the vehicle trajectory. In particular, the heading time

histories in figures 8 and 17 indicate that the piloting

aids, which consisted of the autothrottle and flight director, permitted the maneuver to be performed in a consistent and repeatable manner. A quantitative comparison of

maneuver performance with and without the various

piloting aids was conducted in the second part of the realtime experiment and will be discussed later in this report.

Another metric used to compare the four response

type and HUD combinations was the rms tracking error

of the actual normal load factor from the desired normal

load factor commanded by the guidance algorithm. This

metric provides a relative assessment of how closely a

particular configuration enables the pilot to achieve the

conditions called for by the guidance algorithm. The

pilots were asked to track the load factor director as

closely as possible. Flight director tracking tolerances

were specified as 0.1g for desired performance and 0.2g

for adequate performance as noted in figure 13. Figure 18 again indicates a clear improvement of the LFRC

response type over the RCAH response type. When using

the LFRC system, the pilots experienced an average rms

load factor error of roughly 0.05g, while the RCAH system resulted in an average rms load factor error of

roughly 0.12g. The LFRC system permitted the pilots to

track the load factor guidance cue more precisely,

regardless of the HUD symbology.

A similar comparison of rms tracking error in bank

angle is shown in figure 19. The flight director tracking

tolerances specified in figure 14 for bank angle were 5.0°

for desired performance and 10° for adequate performance. Figure 19 shows that the performance for all four

HUD and response type combinations was well within

the desired tolerance.

Performance Metrics

To compute the dynamic pressure performance metrics, the peak absolute value of the dynamic pressure

error was first determined for each of the real-time runs.

The average peak absolute dynamic pressure error was

then computed for each of the response type and HUD

configurations. The average, maximum, and minimum

values of the peak absolute dynamic pressure error are

shown in figure 20. Error tolerances were specified as

20 lb/ft2 for desired task performance and 30 lb/ft2 for

adequate performance. An analogues plot of the mean

dynamic pressure error is shown in figure 21.

Figure 20 shows that desired performance was

achieved in all runs (recall that autothrottles designed to

regulate dynamic pressure were active in all runs during

this part of the experiment). A slight increase in both the

peak and the mean dynamic pressure errors is apparent

when the RCAH response type is compared with the

LFRC response type. An identical trend in terms of variation in peak and mean absolute altitude errors with

response type and HUD configuration is shown in figures 22 and 23. The altitude error tolerance for desired

performance was 200 ft, and all runs achieved desired

performance. The HUD symbology appeared to have no

appreciable impact on maneuver precision as judged by

the magnitude of altitude and dynamic pressure errors.

Fuel consumption during the maneuver is another

factor to consider when comparing the performance metrics for the four response type and HUD combinations.

Figure 24 shows the variation in fuel consumption with

configuration. On average, the RCAH response type

resulted in slightly greater fuel expenditure during the

maneuver due to increased activity in terms of angle-ofattack variation and associated control surface deflections. This maneuver represents only a small segment of

the entire mission for an air-breathing hypersonic

vehicle, and the relatively small difference in fuel consumption could be significant when integrated over

the duration of the mission. Current air-breathing hypersonic vehicle concepts must emphasize fuel efficiency

throughout the flight envelope to retain any hope of a

achieving a reasonable payload capacity, so the apparent

reduction in fuel consumption provided by the LFRC

system is an attractive feature of this response type.

In summary, part 1 of the real-time experiment indicated that all four pilots preferred the LFRC response

type over the RCAH response type. Pilot comments

emphasized that the load-factor-oriented control system

provided greater harmony with the flight director cue,

which was driven with a load factor command specified

by the guidance system. The impetus for providing a

load-factor-oriented director instead of flight-pathoriented director stems from the hypersonic path-attitude

decoupling effect, which was elaborated on earlier in this

report. The LFRC response type provided the pilots

greater ease and slightly improved precision in control of

the maneuver based on Cooper-Harper ratings, pilot

comments, workload metrics, and task performance metrics. This response type also relieved the pilot of the longitudinal control task once the steady turn was

established thereby producing fewer angle-of-attack

perturbations, less associated control activity, and
12

slightly lower average fuel consumption. Note that this

effect was somewhat influenced by the lack of a trim button in the RCAH implementation and that a different

trend might be observed for other maneuvers that don’t

involve such sustained periods at elevated load factor as

does the cruise turn.

No consistent preference for either HUD was identified because two pilots preferred the conventional

symbology and two preferred the resolver-based symbology. Desired performance specifications were achieved

with both HUD’s, and no significant differences in

maneuver precision or workload were noted. Although

no consistent preference for either HUD symbology

could be identified, each pilot had good justification for

the particular symbology selection. Each of the four test

pilots awarded at least one response type and HUD combination a level-1 flying qualities rating (CHR < 3) for

the cruise turn maneuver and selected a preferred combination for use in part 2 of the real-time experiment.

Discussion of Part 2 Results

The second part of the real-time experiment examined the degradation in task performance and flying qualities ratings as various piloting aids were removed. This

was accomplished by asking the pilots to perform and

compare the maneuver using their preferred flight director and response type under each of the pilot assistance

level (PAL) conditions described earlier in this report.

The PAL setting was varied from 4 (full autopilot task

execution) to 0 (manual control of stick and throttle with

no directors). The intent was to identify the degree of

assistance that must be supplied to the pilot in order to

achieve desired task performance with an acceptable

workload. This section presents the variation with PAL

of pilot workload metrics and of task performance

metrics.

Workload Metrics

The test pilots were asked to assign a Cooper-Harper

flying qualities rating to each of the PAL conditions.

These ratings are shown in figure 25. The PAL 4 runs

involved fully automatic execution of the maneuver with

autopilot and autothrottles, so no Cooper-Harper ratings

were assigned for that condition. Figure 25 shows that

the task was rated as level 1, with an average CHR equal

to 2.6, when PAL 3 was used. The PAL 3 condition corresponds to manual stick control with a flight director

and autothrottle. (See table 5.) When the autothrottle was

deactivated and the throttle director cue was provided

with PAL 2, the average CHR increased to 4. Pilot comments indicate that a considerable workload increase was

associated with tracking the throttle director cue. This

results is due, at least in part, to the fact that the throttle

director calls for a thrust increase during the roll-in phase

of the maneuver, when a relatively high workload is

already required to track the flight director cue. With the

autothrottle deactivated, the pilot is required to roll, pull,

and add power simultaneously during this portion of the

maneuver.

Figure 25 shows that the average CHR awarded was

3.5 when the throttle director cue was removed with

PAL 1. Pilot comments indicated a slight improvement

in the perception of workload when the throttle director

was removed. With PAL 1 the pilots used the digital

readout on the HUD to provide feedback for the manual

regulation of dynamic pressure with throttle. All four

pilots indicated that they were essentially operating the

throttles in an open-loop fashion during the roll-in, relying on their training from earlier observations of the

autothrottle. After the roll-in was completed, they could

check the dynamic pressure readout on the HUD to affect

minor corrections to the throttle setting during the steady

turn, which was part of the maneuver that required much

less attention to the flight director. So, with PAL 1, the

pilots appeared to demote throttle control from a pursuit

task to an open-loop-memorized task during the roll-in,

which could explain the slight reduction in the perceived

difficulty of the task. However, two of the pilots specifically commented that they would prefer the presence of

the throttle director if they were performing an unfamiliar task. All of the pilots commented that, although they

could perform the maneuver satisfactorily with manual

throttle control, they felt that autothrottles should be provided (as with PAL 3) for routine execution of such a

maneuver.

The most notable feature in figure 25 is the dramatic

degradation in acceptability ratings when PAL was

reduced from 1 to 0, which corresponded to removal of

the flight director cue. With PAL 1, the average CooperHarper rating (CHR) assigned by the pilots was 3.5,

whereas with PAL 0 the average increased to 8. This

occurred in spite of the considerable training that the

pilots received in performing the task while using the

higher pilot assistance levels. The pilots indicated that

the task simply could not be performed to adequate specifications on a reliable basis with PAL 0, as had been

noted earlier during the preliminary simulation sessions.

This condition persisted in spite of the change from the

RCAH control system that was used in the preliminary

sessions to the LFRC control system that had been unanimously chosen by the pilots during part 1 of the formal

real-time experiment.

Figure 26 shows the average, maximum, and minimum rms longitudinal stick deflections for the combined

set of real-time runs for all four pilots. The data corresponding to the PAL 4 fully automated runs are included
13

for completeness. The trends observed in figure 26 agree

completely with those in figure 25 and provide an objective confirmation of the assigned pilot ratings. Figure 27

shows a plot of time histories of longitudinal stick deflection from two typical runs conducted by pilot B with

PAL 3 and PAL 0. The pilot was using the LFRC

response type in both runs. A dramatic increase in stick

activity is apparent when the autothrottle and flight director piloting aids are removed, which further corroborates

the trend shown in figure 25.

Performance Metrics

The variation in the task performance metrics with

pilot assistance level generally followed the same trends

that were observed for the workload metrics. Figure 28

shows the average, minimum, and maximum of the peak

absolute value of dynamic pressure error compiled from

the complete set of piloted real-time runs. Bounds that

define desired and adequate task performance are also

indicated in this figure. The PAL 4 fully automatic runs

are included for comparison purposes.

Figure 28 shows that all runs conducted with PAL 3

achieved desired performance specifications, while none

of the runs conducted with PAL 0 achieved even adequate performance specifications. Figure 29 shows the

analogous plot for the peak absolute altitude error. Once

again, none of the runs conducted with PAL 0 achieved

even adequate performance. The variation of the mean

absolute dynamic pressure error with PAL is shown in

figure 30, and mean absolute altitude error is shown in

figure 31. The mean errors shown in these plots follow

the same trends observed in the peak error plots in figures 28 and 29. A significant increase in altitude error is

evident when the maneuver is conducted with manual

stick control and autothrottles (PAL 3) instead of the full

autopilot (PAL 4), but desired task performance is still

achieved. No significant variation is observed in task

performance with PAL’s 3, 2, and 1. Although the pilot

comments reflected a preference for autothrottles, only

minor differences were apparent in the task performance

metrics.

Time histories from typical runs with the LFRC

response type and PAL 0 for each of the four pilots are

plotted in figure 32. The time histories show that none of

the PAL 0 cases satisfied the adequate performance tolerances. Apparently the failure of the pilot to meet

dynamic pressure task performance specifications with

PAL 0 is related to a basic inability of the pilot to adequately regulate the vehicle vertical flight path through

the turn. The poor task performance is due to the same

path-attitude decoupling phenomenon that was noted in

the preliminary simulation sessions, although a pitchrate-command–attitude-hold (RCAH) system was used

in those sessions. Equation (3) illustrated that even minor

vertical flight path perturbations can yield extreme

dynamic pressure perturbations at hypersonic speeds.

Figure 32 shows that the pilots were attempting to

counter the resulting altitude and dynamic pressure perturbations with longitudinal stick (normal load factor)

and throttle inputs, which lead to a very high perception

of workload and poor task performance. The bank angle

and heading time histories illustrate the lack of precision

and repeatability that is experienced when the maneuver

is conducted without the aid of a flight director. Excessive angle-of-attack perturbations also resulted from the

frequent longitudinal inputs. Note that the pilots had an

opportunity to observe the fully automatic execution of

the maneuver with PAL 4 in the simulator prior to conducting the maneuver themselves with PAL 0.

Figure 33 presents time history plots of typical

maneuvers executed by pilot D using each of the pilot

assistance levels with his preferred HUD and response.

The PAL 4 fully automatic execution is included for

comparison. The flight director guidance supplied in the

PAL settings of 1–3 permitted the pilot to effectively

cope with the high-speed path-attitude decoupling and to

achieve desired performance as indicated by the dynamic

pressure and altitude time histories. Recall that the flight

director cue (regardless of the pilot’s preferred symbology) is driven by the load factor command specified by

the guidance algorithm rather than by a vertical flight

path error. Examination of the altitude, bank angle, and

load factor time histories for PAL’s 1–4 in figure 33

shows that the problem of vertical flight path control is

solved by the flight director guidance strategy of minor

bank angle adjustments to regulate altitude through the

turn, while a normal load factor of 2g is maintained. A

significant reduction in angle-of-attack perturbations is

realized when compared with the PAL 0 (no flight director) case. Throttles are used primarily to counter the drag

increment produced by the load factor increase from 1g

to 2g. This is in contrast to the pilot’s intuitive strategy of

modulating load factor to regulate altitude through the

turn while using throttle to null dynamic pressure perturbations as the PAL 0 traces show. A comparison of the

heading time histories in figure 33 with those in figure 32

indicates the improvement in task repeatability that is

gained through the use of the flight director with

PAL’s 1–4.

The variation in fuel required to complete the

maneuver is plotted versus PAL in figure 34. The figure

shows that dramatic fuel savings is realized as a result of

following the guidance strategy described previously. In

fact, the average fuel savings for this maneuver, which

lasts approximately 90 sec, is greater than 1000 lb and is

a substantial fraction of the payload capacity for such a

configuration.
14

A more subtle trend is also apparent when comparing fuel consumption of PAL 4 (full autopilot and autothrottle) with that of PAL 2 (manual throttle with throttle

director). The maneuvers executed with PAL 2 consistently used about 40 lb less fuel than those executed with

PAL 4. Because fuel efficiency is a key factor in the feasibility of the hypersonic air-breathing vehicle concept,

the answer to why the pilots were able to manually

achieve slightly better efficiency than the autothrottles

while still achieving desired performance is important. A

comparison of the mean dynamic pressure errors in figures 30 and 34 reveals the reason. With the larger mean

dynamic pressure excursions permitted through the

maneuver, which did not violate the task performance

specifications, the pilots achieved desired performance

and realized a fuel savings. This practice is illustrated by

the dynamic pressure time history for the PAL 2 case in

figure 33, which shows a dynamic pressure deficiency of

approximately 8 lb/ft2 throughout the steady segment of

the maneuver. The propulsive efficiency of the engine

model used in this investigation was not significantly

penalized in response to off-nominal dynamic pressure

conditions, so the pilots’ technique resulted in a slight

fuel savings. The actual sensitivity of propulsive performance to flight condition for this type of engine is highly

uncertain. Given an accurate propulsive model, a fueloptimal guidance algorithm can be envisioned which

would direct the pilot along the most efficient trajectory

subject to propulsive performance and aerodynamic heating constraints. This topic is currently the subject of considerable research. The guidance algorithm described in

reference 13 includes no such considerations, and the

autothrottles merely sought to regulate dynamic pressure

at the prescribed reference value of 2000 lb/ft2. The

observation emphasizes the importance of fuel-optimal

trajectory determination for this type of configuration.

The analysis of pilot workload metrics and task performance metrics from part 2 of the real-time experiment

indicates which aids must be provided to the pilot to

achieve desired task performance with an acceptable

workload. The complete set of piloting aids, which consist of the autothrottle, throttle director, and flight director, improves the average of Cooper-Harper ratings from

8 to 2.6. Level-1 flying qualities ratings (CHR < 3) and

desired task performance were achieved when the

maneuver was performed with PAL 3, that is, with the

flight director and autothrottle provided to the pilot.

Level-2 flying qualities ratings (4 < CHR < 6) and

desired task performance were achieved when the flight

director was provided and the throttles were under manual control (PAL 1), as long as the pilot had a digital

readout of dynamic pressure and had received sufficient

training. In the absence of the autothrottle, throttle director, and flight director cues (PAL 0), the flying qualities

were rated level 3 (7 < CHR < 9) or worse, and achievement of even adequate task performance was never possible. These results suggest that a flight director driven

by a guidance algorithm tailored to the specific characteristics of high-speed flight should be considered a

flight-critical element of the control system for an airbreathing hypersonic vehicle. Autothrottles should be

provided to reduce pilot workload and should probably

be treated as flight-critical control elements based on

pilot comments. When sufficient training was provided,

the pilot was able to maintain adequate control of the

vehicle trajectory during the cruise turn in the absence of

the autothrottle, but this is contingent upon the throttleability of scramjet propulsion systems at hypersonic

speeds and their operational sensitivity to flight condition

perturbations.

Concluding Remarks

An investigation has been conducted to develop and

evaluate a series of piloting aids for maneuver control of

an air-breathing hypersonic vehicle. Stringent propulsion

constraints and nonintuitive high-speed flight effects

associated with maneuvering in the hypersonic regime

raise the question of whether manual control of such a

vehicle should even be considered.

The objectives of this research were to determine the

extent of manual control that is desirable for a vehicle

maneuvering in this regime and to identify the form of

aids that must be supplied to the pilot to make such control feasible. A piloted real-time motion-based simulation of an air-breathing hypersonic vehicle concept was

used for this study, and the investigation focused on a

single representative cruise turn maneuver at a Mach

number of 7.86 and an altitude of 85 000 ft.

Several maneuver control aids were developed as a

result of pilot comments and control problems encountered during a series of preliminary maneuver simulation

sessions. These aids consisted of an autothrottle, throttle

director, autopilot, flight director, and two head-up display (HUD) configurations. The first of the two HUD

configurations contained relatively conventional flight

director symbology elements. The second HUD presented a graphical representation of the vehicle lift vector

in terms of normal load factor and bank angle. The algorithm, which was used to drive both HUD configurations,

was based on a normal load factor specified by the guidance law. The impetus for providing the load-factororiented guidance instead of flight-path-oriented guidance stems from the hypersonic path-attitude decoupling

effect. This effect introduces a significant lag into the

vehicle’s flight path response to attitude changes, which

makes direct manual control of flight path impractical.

Two longitudinal control response types consisting of a
15

rate-command–attitude-hold (RCAH) system and a loadfactor-rate–load-factor-hold (LFRC) system were also

developed.

A formal real-time experiment was then conducted

which included two parts and the participation of four

NASA test pilots. The first part of the experiment identified the pilots’ preferences from a combined set of the

two response types and the two HUD configurations to

determine which combination provided the greatest ease

and precision of control. The second part of the experiment examined the degradation in task performance and

pilot satisfaction as various aids were removed from the

pilot’s use; these results were compared with maneuver

performance in the fully automatic mode.

The first part of the real-time experiment revealed

that all four pilots preferred the LFRC response type over

the RCAH response type. This preference was due in part

to the greater harmony of the load-factor-oriented control

system with the flight director cue, which was driven

with a load factor command specified by the guidance

system. This response type also provided pilot relief

from the longitudinal control task after the steady turn

was established, which thereby produced fewer angle-ofattack perturbations, less associated control activity, and

slightly lower fuel consumption. Note that this effect was

somewhat influenced by the lack of a trim button in the

RCAH implementation and that a different trend might

be observed for other maneuvers that do not involve such

sustained periods at elevated load factor as does the

cruise turn. The LFRC response type provided pilots

greater ease and slightly improved precision in control of

the maneuver on the basis of pilot comments, CooperHarper flying qualities ratings, workload metrics, and

task performance metrics.

No consistent preference for either HUD was identified because pilot opinions were evenly split between

the two symbologies. Desired performance specifications

were achieved with both HUD’s, and no significant differences in maneuver precision or workload were noted.

Each of the four test pilots awarded at least one response

type and HUD combination a level-1 flying qualities rating (CHR < 3) for the cruise turn maneuver and selected

a preferred combination for use in the second part of the

real-time experiment.

The second part of the experiment indicated which

aids must be provided to the pilot to achieve desired task

performance with an acceptable workload. The complete

set of piloting aids, which consisted of the autothrottle,

throttle director, and flight director, improved the average of Cooper-Harper ratings from 8 to 2.6. The flight

director was the most critical of these aids, and the cruise

turn maneuver became essentially unachievable to adequate performance specifications in the absence of this

flight director. Level-1 flying qualities ratings and

desired task performance were achieved when the flight

director and autothrottle were provided to the pilot.

Level-2 flying qualities ratings and desired task performance were achieved when the flight director was provided and the throttles were under manual control, as

long as the pilot had a digital readout of dynamic pressure. In the absence of the autothrottle, throttle director,

and flight director aids, the flying qualities were rated

level 3 or worse, and achievement of even adequate task

performance was never possible.

Hypersonic maneuver control will be significantly

complicated by high-speed flight effects such as pathattitude decoupling and centrifugal relief. Propulsion

system sensitivity to flight condition perturbations will

present additional challenges to the pilot of such a vehicle. The severity with which these nonintuitive effects

alter the pilot’s ability to control the vehicle makes some

form of assistance imperative. This investigation focused

on the execution of a single representative cruise turn

maneuver at a Mach number of 7.86 and an altitude of

85 000 ft. The results indicate that a flight director driven

by a guidance algorithm tailored to the specific characteristics of high-speed flight should be considered a

flight-critical element of the control system for an airbreathing hypersonic vehicle. Autothrottles should be

provided to reduce pilot workload and, on the basis of

pilot comments, be treated as a flight-critical control element for an air-breathing hypersonic vehicle.

NASA Langley Research Center

Hampton, VA 23681-0001

July 14, 1995
16

Appendix A

Propulsion Model

The hypersonic propulsion model used in this simulation was based primarily on the propulsive database

provided in reference 15. However, the propulsive database supplied by reference 15 contains several somewhat

artificial and unrealistic features required for operation of

the Program to Optimize Simulated Trajectories (POST).

For example, the winged cone propulsive database

includes data points at fuel equivalence ratios as high as

100 to prevent POST from extrapolating to negative specific impulses in the lower fuel equivalence ratio regions.

Such high fuel equivalence ratios are unrealistic and

were not included in the Langley Visual/Motion Simulator (VMS) hypersonic propulsion model. A modified

engine model was developed that was based on only the

most reliable data presented in reference 15. This appendix describes the characteristics of the modified propulsion model.

Symbols

interpolation table function number 1

interpolation table function number 2

ISP specific impulse, sec

specific impulse for equivalence ratio of 1, sec

M Mach number

dynamic pressure, lb/ft2

s Laplace transform parameter

T axial thrust, lb

V inertial velocity, ft/sec

WF fuel flow rate, lb/sec

WF|η=1 fuel flow rate for equivalence ratio of 1, lb/sec

η fuel equivalence ratio

τ

eng engine response time constant

Subscripts:

cmd commanded value

max maximum allowable value

Model Description and Implementation

This modified propulsion model includes a suitable

input for use by the autothrottle control law. A penalty is

applied for fuel equivalence ratios η > 1.0; equivalence

ratio information is included up to 10.0, although excursions as high as 10.0 would be extremely inefficient and

improbable. A flowchart of the modified engine model is

shown in figure A1.

The original propulsive database presented in reference 15 also included a weak dependence on dynamic

pressure, which forced the interpolations for specific

impulse and thrust coefficient to be three-dimensional.

The dynamic pressure dependence has been eliminated

from the interpolations because this was a weak and

somewhat uncertain effect, although dynamic pressure

remains in the model as a multiplier of thrust coefficient.

The revised propulsion model yields a net thrust

which acts along the vehicle body X-axis. The effects of

angle of attack and sideslip, body angular rates, and control surface deflections on thrust and specific impulse are

not accounted for in this model. In this simulation, fuel

flow rate can be commanded by the pilot or by an autothrottle feedback control law; the choice depends on the

mode selected. This fuel flow rate command is then used

with Mach number and dynamic pressure to compute a

fuel equivalence ratio. The equivalence ratio is then used

in a one-dimensional interpolation table to compute a

specific impulse that is valid for equivalence ratios ≤1.

A penalty is then applied for equivalence ratios >1 in the

form of an upper limit on specific impulse. This specific

impulse is then multiplied by fuel flow rate to determine

the resultant thrust. The procedure for calculation of

equivalence ratio, specific impulse, and thrust from commanded fuel flow rate is as follows:

1. Determine fuel flow rate WF from commanded fuel

flow rate by simulation of engine fuel flow

dynamics (a value of τ

eng

 = 0.5 was used in the

simulation)

2. Use natural log of Mach number in figure A2 to

interpolate for the quantity f1(M), which is a onedimensional interpolation.

3. Determine ideal fuel flow rate WF|η = 1 from f1(M) as

follows:

where exp is the inverse of the natural log.

4. Determine fuel equivalence ratio η as follows:

5. Use Mach number in figure A3 to interpolate for the

specific impulse achieved by this propulsion system when the fuel equivalence ratio η = 1.

6. Use the fuel equivalence ratio with figure A4 to interpolate for the specific impulse limit penalty below and

f 1

f 2

ISP′

q

WFcmd

WF

WFcmd

τengs + 1 = ---------------------

WF η = 1 = q exp[ ] f 1( ) M

η

WF

WF η = 1

= ---------------------

ISP′
17

apply limit as shown in the block diagram in figure A1

to determine actual specific impulse ISP.

7. Multiply specific impulse ISP by fuel flow rate WF to

determine axial thrust T.

An explanation of steps (5) and (6) follows. Specific

impulse is a measure of fuel efficiency in an airbreathing engine. For hypersonic propulsion systems, the

best efficiency occurs at a fuel equivalence ratio of 1.

Figure A3 (step 5) shows the specific impulse that is

achieved by this propulsion system at various Mach

numbers when the fuel equivalence ratio is set to 1. As

equivalence ratio is adjusted above 1, the fuel efficiency

drops off (i.e., specific impulse decreases). Step (6)

applies a correction factor to the specific impulse from

step (5) to reflect this decrease in efficiency. Figures A2–

A4 are based on the most realistic segments of the original winged-cone database provided in reference 15.

Interpolation data are included in figures A2–A4.

Two limiters are included in the block diagram

shown in figure A1 to ensure that Mach number and fuel

equivalence ratio remain within the limits of this model.

This model is appropriate for Mach numbers between 0.3

and 25 and fuel equivalence ratios between 0.1 and 10.

ISPmax = f 2( ) M /η
18

Figure A1. Block diagram of propulsion model.

Figure A2. Interpolation data for function f1(M).

e Mach x ×

WF

ln η

0.3

25

Limiter

Figure A2 

interpolation

/

1

τeng

WF s + 1 cmd

Fuel flow

dynamics

η

ISP

×

WF

T

Mach

ISP

Mach

Figure A3 

interpolation

'

0.1

10

Mach

Figure A4 

interpolation

/

Limiter

Limiter

ISPmax

q

f 2(M)

f 1(M) WF|η =1

ISPmax

.14207

–1.61141

–3.16976

–1.20000 –.15065 3.20000

A

AAAA

AAAAAAAAA

A

A

A

A

A

A

A

A

A

A

A

A

A

ln 

f 1(M)
19

Figure A3. Specific impulse function data.

Figure A4. Interpolation data for function .

AAA

A

A

A

A

A

A

3600

1300

0 .90

1

0.000077 + 0.00004688 M

Mach number

3100 AA

A

A

A

A

A

A

1.50A

A

A

A

A

A

A

2.00 4.28 A

A

A

A

A

A

A

ISP '

0

AA

AAAA

A

A

A

A

A

A

A

A

A

A

A

A

A

3118

4089

2530

4 8 25 A

A

A

A

Mach number

2868

AAAAAAAAAA

f 2(M)

f 2( ) M
20

Appendix B

Description of Inner-Loop Flight Control Laws

This appendix describes the derivation of the innerloop flight control laws, which stabilize the vehicle

dynamics, track load factor and bank angle commands

from the pilot or autopilot, and maintain zero sideslip

during maneuvers. A simple pole-placement technique is

used, which allows specification of the inner-loop

responses in terms of desired frequency and damping.

Figure 12 shows a flowchart of the real-time simulation,

which depicts the relation of the inner-loop flight controls in relation to other components of the simulation.

Three separate inner-loop control laws are synthesized to

drive the symmetric elevon, differential elevon, and rudder deflections for pitch, roll, and yaw control. The estimation of sideslip angle with a lateral accelerometer and

the concept of a complementary filter to reduce highfrequency control action resulting from atmospheric disturbances are also described.

Symbols

AY lateral acceleration in body-axis coordinate

system, ft/sec2

b wing span, ft

g acceleration due to gravity, ft/sec2

I

X rolling moment of inertia, slug-ft2

I

XZ product of inertia, slug-ft2

I

Z yawing moment of inertia, slug-ft2

K1 effective pitch loop proportional gain

K2 effective pitch loop integral gain

sideslip feedback gain to aileron

rudder deflection crossfeed gain to aileron

bank angle feedback gain to aileron

roll rate feedback gain to aileron

Kq pitch rate feedback gain to elevons

sideslip feedback gain to rudder

yaw rate feedback gain to rudder

KsE1 longitudinal stick gain for RCAH system

KsE2 longitudinal stick gain for LFRC system

L, M, N rolling, pitching, and yawing moments, ft-lb

m mass, slug

NV normal load factor with respect to velocity

vector, g units

n normal load factor, g units

p roll rate in body-axis coordinate system,

deg/sec

q pitch rate in body-axis coordinate system,

deg/sec

dynamic pressure, lb/ft2

r yaw rate in body-axis coordinate system,

deg/sec

s Laplace transform parameter

u, v, w body-axis velocities, ft/sec

V inertial velocity, ft/sec

W weight, lb

xb

, yb, z

b vehicle body axes

x actuator feedback signal

X, Y, Z body-axis force components, lb

α angle of attack, deg

β sideslip angle, deg

∆ characteristic equation

δa aileron deflection, positive right trailing edge

down, deg

δe elevon deflection, positive trailing edge

down, deg

δr rudder deflection, deg

ζOL open-loop damping ratio (pitch)

ζa aileron control loop damping ratio

ζr rudder control loop damping ratio

ζsp short-period damping ratio

ζ0 damping ratio of introduced actuator and filter

dipole

θ pitch angle, deg

λ distance from c.g. to lateral accelerometer,

positive forward, ft

τL actuator feedback filter time constant

τ

a actuator lag time constant

τ

β sideslip complementary filter time constant

effective pitch loop filter time constant

φ roll angle, deg

ωOL open-loop natural frequency (pitch), rad/sec

ωa aileron control loop natural frequency, rad/sec

ωr rudder control loop natural frequency, rad/sec

ω

sp short-period natural frequency, rad/sec

Kaβ

Kaδr

Kaφ

Kaφ ̇

Krβ

Krr

q

τ′
21

ω0 frequency of introduced actuator and filter

dipole, rad/sec

Superscripts:

. first derivative with respect to time

.. second derivative with respect to time

^ estimated value

~ complementary filtered value

Subscripts:

acc accelerometer signal

cmd commanded value

err error signal

RW relative to wind

0 nominal or trim value

Pitch Control1

The primary element of the longitudinal control

architecture is a pitch rate command loop. The object is

to control pitch rate for a plant with dynamics that are

approximated by the second-order system given in

equation (B1).

(B1)

The basic pitch rate command architecture is shown

in figure B1. The pitch rate error is multiplied by gain

Kq and then fed into a first-order actuator model with

time constant τa. Equation (B2) expresses the control

deflection-to-pitch rate error transfer function for the network shown in figure B1. This transfer function can be

manipulated to achieve equations (B3)–(B6), which

illustrate that the addition of the pitch control filter loop

around the actuator shown in figure B1 effectively produces a proportional-integral control action on the pitch

rate error signal.

(B2)

(B3)

1Adapted from unpublished notes by Frederick J. Lallman, “Hypersonic Pitch Control Laws,” April 19, 1990.

where

(B4)

(B5)

(B6)

The dynamics of the closed-loop system are

described by equation (B7) and the associated characteristic equation (B8).

(B7)

(B8)

where

(B9)

(B10)

(B11)

The standard form of the fourth-order characteristic

equation for the short-period pitch mode and the introduced actuator and filter quadratic is given by equation (B12), and the associated closed-loop pitch rate

transfer function is given by equation (B13).

(B12)

(B13)

w ̇

q ̇

Zw V

Mw Mq

w

q       Zδe

Mδe

δe = +

δe

qerr

--------( )s

Kp

1

τas + 1 ----------------    

1 1

τas + 1 ----------------     1

τLs + 1 -----------------     –

= ------------------------------------------------------

δe

qerr

--------( )s K1

K2

s

+ ------     1

τ′s + 1 ----------------     =

K1 Kp = τL

τa τL + ----------------

     

K2 Kp

1

τa τL + ----------------     =

τ′

τaτL

τa τL + = ----------------

w ̇

q ̇

δe ̇

x ̇

Zw V Zδe 0

Mw Mq Mδe 0

0

Kp

τa

------ 1

τa

–---- 1

τa

----

0 0 1

τL

----- 1

τL

–-----

w

q

δe

x               0

0

Kp

τa

------

0

qcmd = –

∆ s s

1

τa

---- 1

τL

+ + -----     s

2 2ζOLωOLs ωOL

2 ( ) + + =

KpMδe

τa

------------------ s

1

τL

+ -----     s

1

t

θ2

+ ------     –

ωOL

2 ZwMq V Mw = –

2ζOLωOL Zw Mq = –( ) +

1

t

θ2

------

Zδe

Mδe

---------Mw Zw = –

∆ s

2 2ζspωsps ωsp

2 ( ) + + s

2 2ζ0ω0s ω0

2 = ( ) + +

q

qcmd

----------( )s

KpMδe

τa

---------------- s 1

τL

+ -----     s 1

tθ2

+ ------    

s

2 2ζspωsps ωsp

2 ( ) + + s

2 2ζ0ω0s ω0

2 ( ) + +

= ----------------------------------------------------------------------------------------------------
22

Values for the pitch rate error gain Kq and the filter

time constant τL may be obtained by prescribing desired

values for the closed-loop short-period characteristics

and equating coefficients of like powers in equations (B8) and (B12).

A simple modification of the architecture shown in

figure B1 will eliminate the zero in the pitch rate transfer

function at 1/τL. The elimination of this zero will provide

the system with more desirable transient response characteristics. This is achieved by moving the pitch rate

command signal from driving the actuator directly

through gain Kq to a position where it enters the actuator

through the control filter and a duplicate of the gain Kq.

This modification is shown in figure B2. The resulting

pitch rate transfer function is shown in equation (B14).

(B14)

Other elements were added to the pitch control

architecture to provide the pitch-rate-command–attitudehold and load-factor-rate–load-factor-hold response

types. These additions are shown in the complete pitch

control block diagram in figure B3. The pitch rate command input to the system shown in this figure was generated by one of three sources, which depends upon the

response type and pilot assistance level selected. First,

figure B3 shows that if the autopilot was not active

(PAL < 4) and the response type selected was RCAH, the

pitch rate command was generated by the pilot’s longitudinal stick deflection through a scaling gain KsE1. A

proportional-integral (PI) network on pitch angle error

was used to provide the attitude-hold control action. A

detent region of ±0.2 deg/sec was placed on the pilot’s

pitch rate command signal to delimit the operation of this

pitch attitude-hold feature. When the pitch rate command

produced by the pilot’s stick deflection was outside the

detent region, the reference pitch attitude θ0

 was set

equal to the actual pitch attitude, so that the error signal

being fed through the attitude-hold PI network was zero.

When the stick entered the detent region, the pitch attitude was sampled, and this value was used as the pitch

attitude reference for the PI network. The gains in the PI

attitude-hold network were tuned to provide good

response in turbulence, and the integrator was limited at

10°. A delay of 0.5 sec between the instant the stick

enters the detent and the instant the reference pitch attitude is sampled was necessary to prevent an excessively

abrupt arrest of the pitch rate with an associated overshoot of the reference value.

The second possible source of the pitch rate command signal shown in figure B3 was selected if the autopilot was not active and the designated response type was

load factor rate command (LFRC) hold. In this case, the

longitudinal stick deflection was fed through a scaling

gain KsE2 and integrated to produce a load factor command, which was then differenced from an estimate of

the actual load factor based on accelerometer signals and

angle of attack. This load factor error was then fed

through a PI network and converted into a pitch rate

command by equation (B15). Gains in the PI network

were tuned according to pilot preference during the

preliminary simulation sessions. Limits were placed on

the load factor command to prevent the signal from

exceeding 2g.

(B15)

The final source of the pitch rate command shown in

figure B3 was selected if the autopilot was activated. In

this case, the load factor command from the guidance

law described in reference 13 was differenced from the

estimated load factor signal and fed through the same PI

network used in the LFRC response type. Because the

same signal that had been used to drive the flight director

while the vehicle was in manual control was used to

drive the autopilot, the pilots felt that the automatic systems performed in a predictable and familiar fashion.

Yaw Control

The rudder is used to maintain zero sideslip during a

maneuver. Figure B4 shows a block diagram of the yaw

control law. The yaw rate and sideslip equations are

given by

(B16)

(B17)

The final term in equation (B17) involving gravity

will be negligible in hypersonic flight because the velocity is very high. Also, the effect of rotary derivatives is

often small at hypersonic speeds, because velocity is in

the denominator of the dimensionalizing expression.

Neglection of the rotary derivative and inertial coupling

terms in equations (B16) and (B17) and a Laplace transformation yield the following dynamic system:

(B18)

(B19)

q

qcmd

----------( )s

KpMδe

τa

---------------- s 1

tθ2

+ ------    

s

2 2ζspωsps ωsp

2 ( ) + + s

2 2ζ0ω0s ω0

2 ( ) + +

= ----------------------------------------------------------------------------------------------------

qcmd NV err

g

V

---     180

π

--------     1

Zw

------     = –

r ̇ Nββ Np p Nrr Nδaδa Nδrδr

IXZ

IZ

= + + + + + -------- p ̇

β

 ̇ Yββ Y p p Yrr Yδa

δa Yδrδr = + ++ + – r

+ g

V

--- cosθ0 sinφ

sr s( ) Nββ( )s Nδa

δa( )s Nδrδr = + + ( )s

sβ( )s Yββ( )s Yδa

δa( )s Yδrδr = + + ( )s – r s( )
23

A rudder deflection command is defined with yaw rate

and angle-of-sideslip feedback in the following form:

(B20)

Substitution of the rudder deflection command into equations (B18) and (B19) and neglection of yaw and side

forces due to aileron deflection yield

(B21)

(B22)

This system can be expressed in matrix form as

(B23)

This matrix yields the following second-order characteristic equation for the closed-loop system:

(B24)

 The coefficients of this characteristic equation are

expressed in equations (B25) and (B26) in terms of frequency ωr and damping ζr

 for the second-order system.

(B25)

(B26)

Neglection of the relatively small terms and solution of

the feedback gains Krβ and in terms of frequency

and damping yield

(B27a)

(B27b)

The feedback gains required to produce a given

closed-loop frequency and damping combination can

now be quickly determined. This permits the designer to

place the closed-loop poles at a location which might be

desirable from a handling qualities perspective.

Angle-of-Sideslip Estimation

The reliability of air data measurements during some

phases of hypersonic flight may be uncertain. The control system designed in this report will utilize an angleof-sideslip feedback that is constructed from a lateral

accelerometer reading rather than an air data sensor

array. The lateral accelerometer will be placed at the

instantaneous center of rotation in response to a rudder

deflection, which eliminates the need to correct the

accelerometer reading for rudder effects. In actual practice, the signal needs to be filtered appropriately to eliminate corruption of the measurement due to structural

vibrations. The equation for the lateral acceleration measured by an accelerometer located a distance λ forward

of the c.g. is given by

(B28)

The solution of this expression for β yields an angle-ofsideslip measurement in terms of the accelerometer reading and the rudder deflection

(B29)

By placing the accelerometer at the lateral center of rotation, the influence of rudder deflection on the angle-ofsideslip measurement is eliminated. Selection of λ so that

(B30a)

yields

(B30b)

This expression for sideslip can be substituted into

equation (B20) to make the lateral inner-loop controller

independent of air data measurements. Note that the

expression for λ depends on parameters which change

with flight condition, i.e., the lateral center of rotation is

not constant over the range of hypersonic flight conditions which the vehicle will experience. Because relocation of the accelerometer as flight condition varies is

unreasonable, the instrument should probably be placed

at some location that represents the average position of

the center of rotation for those flight conditions at which

the air data system measurements are the least reliable.

For the simulated time histories presented in this report,

which include flight conditions at Mach 6, 10, 15, and

20, the lateral accelerometer was placed at the center of

rotation for the Mach 6 flight condition.

δr( )s Krββ( )s Krr = + r s( )

sr s( ) Nββ( )s Nδr Krββ( )s Krr + r s( ) – 0 – [ ] =

sβ( )s r s( ) Yββ( )s Yδr Krββ( )s Krr + r s( ) + 0 – – [ ] =

s NδrKrr – Nβ – NδrKrβ –

1 Yδr

Krr – s Yβ – YδrKrβ –

r s( )

β( )s       0

0 =

s

2 Yβ Yδr

Krβ NδrKrr – ( ) + + s

+ Nβ Nδr

Krβ Krr Nδr

Yβ YδrNβ [ ] + + ( ) – = 0

Yβ Yδr

Krβ NδrKrr – 2 ( ) + + ζrωr =

Nβ Nδr

Krβ Krr NδrYβ YδrNβ + + ( ) – ωr

2 =

Krr

Krβ

ωr

2 Nβ –

Nδr

= -------------------

Krr

2ζrωr Yβ YδrKrβ + +

Nδr – = ----------------------------------------------------

AY acc V0 Yββ Yδr

δr +     λ Nββ Nδrδr +     = +

β

ˆ

AY acc δr V0Yδr λNδr +     –

V0Yβ λNβ + = ----------------------------------------------------------------

λ

V0Yδr –

Nδr

= ------------------

β

ˆ AY acc

V0Yβ λNβ + = ------------------------------
24

Complementary Filter

Sudden crosswind gusts would be fed directly into

the lateral control surface deflections by sideslip feedback under the rudder control law shown in equation (B20). To avoid sharp peaks in the commanded position of control surfaces, a low-pass filter is added to the

sideslip angle feedback loop. The filter is complemented

with derivative feedback to produce a unity transfer function, which eliminates the filter effect on the stability of

the closed-loop system and prevents it from interfering

with the response to pilot or guidance system commands.

Because of the derivative feedback through the filter, this

network is referred to as a “complementary filter.” The

expression for the filtered angle-of-sideslip feedback is

(B31)

where is obtained from equation (B29), and is given

by

(B32)

Rotary derivatives have been neglected in equation (B32). The time constant can be selected to provide

a reasonable amount of disturbance rejection and a favorable response characteristic. Equation (B31) can be substituted for sideslip feedback in the rudder control law

equation (B20). Ideally, the complementary filter has a

unity transfer function and does not affect the stability of

the closed-loop system. In actual practice, the degree to

which the transfer function of equation (B31) differs

from unity depends on the accuracy of the estimated

derivative signal used in this equation. To this extent, the

complementary filter acts only to reduce high-frequency

control action in response to atmospheric disturbance

inputs but not to commanded control inputs.

Roll Control

Roll control is addressed with the use of ailerons as

the primary control effector to track bank angle commands issued by the pilot or autopilot. Figure B5 shows a

block diagram of the basic roll control law. The roll rate

and bank angle equations are given by

(B33)

(B34)

Neglection of the rotary derivative and inertial coupling

terms in equation (B33) yields

(B35)

The simplifying assumption is made that in equation (B34). A feedback control law for aileron deflection

is now defined, which will permit decoupling of yawing

and rolling motions by elimination of the effect of rudder

deflection and angle of sideslip in equation (B35).

(B36)

In this expression, φerr is defined as the commanded bank

angle minus the actual bank angle. This control law also

includes roll attitude and rate feedback and provides

a means to specify frequency and damping in roll

response. The angle-of-sideslip signal used in this

expression will be the lateral accelerometer complementary filtered estimate in equation (B31). Note that the

rudder deflection command is also used in this control

law. Substitution of the aileron control law into equation (B35) yields

(B37)

To achieve the desired decoupling of roll and yaw

responses, define

(B38a)

(B38b)

which, upon substitution into equation (B37), produces

the expression

(B39)

A Laplace transform of equation (B39) yields the following second-order characteristic equation:

(B40)

β

 ̃ τββ

 ̇

^

τβs + 1 ----------------- β

ˆ

τβs + 1 = + -----------------

β

ˆ β

 ̇

^

β

 ̇

^ Yββ

ˆ Yδa

δa Yδrδr = + + – r

p ̇ Lββ Lp p Lrr Lδa

δa Lδrδr

IXZ

IX

= + ++ + + -------- r ̇

φ ̇ p r tan θ0 = + cosφ

p ̇ Lββ Lδa

δa Lδrδr = + +

φ ̇ ≈ p

δa Kaβ

β Kaδrδr Kaφφerr Kaφ ̇ = ++ + φ ̇

 ̇ ̇φ Lββ Lδr

δr Lδa(Kaββ Kaδrδr = + + +

+ Kaφ

φerr Kaφ ̇φ ̇ + )

Kaβ

Lβ –

Lδa

= ---------

Kaδr

Lδr –

Lδa

= ----------

 ̇ ̇φ Lδa

Kaφφerr LδaKaφ ̇ = + φ ̇

s

2 Lδa

Kaφ ̇s LδaKaφ– 0 – =
25

The feedback gains in this expression are easily solved

for in terms of desired frequency and damping of the

closed-loop system to yield

(B41a)

(B41b)

These gains, along with those in equations (B38), are

used in equation (B34) to produce the final aileron

control law. The rotary terms neglected in the derivation

of this control law were of little significance at the hypersonic flight conditions which were simulated. The inherent aerodynamic roll damping of the vehicle is negligible

at these flight conditions, so most of the damping is supplied by the roll rate feedback loop in the aileron control

law.

Kaφ

ωa

2

Lδa – = ----------

Kaφ ̇

2ζaωa

Lδa – = ---------------
26

Figure B1. Basic architecture of pitch rate control loop.

Figure B2. Diagram of modified pitch rate control loop.

Figure B3. Pitch control law used in real-time simulation.

Kq

+

–

+

+

1

τa s + 1

1

τL s + 1

q

qcmd

qerr δe

x

Kq

+

–

+

+

1

τa s + 1

1

τL s + 1

q

qcmd

δe

x

Kq

Autopilot engaged:

• AP switch in position 2

• Autopilot drives inner-loop 

 controls directly

 

Autopilot disengaged:

• AP switch in position 1

• Pilot uses side arm controller 

 in RCAH or LFRC mode

Control modes:

• Rate command-attitude hold

 RCAH = 1, LFCH = 0

• Load factor command hold

 RCAH = 0, LFCH = 1

qcmd(pilot)

+

–

+

+

1

s KnI +4

-4

Limited 

integration

Kn

NVcmd

AP switch

1

2

VRW, ft/sec

AP switch

1

2 Function

F1

ax, ft/sec2

–

+

Actuator

1

τa s + 1 δe

+

+

q

1

 F1: NVest = –az cos α + ax sin α

g

+

Longitudinal

stick deflection

Ks E1 RCAH

+

–

θ

θ0

θerr

KθI

1

s

Kθ

+

32.17 /

Nmax

Nmin

Limiter

LFCH

+ + 1

s

Ncmax

Limited 

integration

Ncmin

NCR

NVcmd(pilot)

NVcmd(autopilot)

qcmd

Kq

+

+

Kθ

– 180

π Zw NVest

NVerr

Ks E2

τL s + 1

g, ft/sec2

α, deg

az, ft/sec2
27

Figure B4. Yaw control law used in real-time simulation.

Figure B5. Roll control law used in real-time simulation.

Krr

Krβ

r

Actuator

1

τa s + 1 δr

+

+

+

+

β

β

1

τβ s + 1

β

~

τβ

φcmd

φ

–

+

δr

Actuator

1

τa s + 1

δa

+

+

+

p

–

+

Ksa Lateral stick 

deflection

pcmd

AP switch

1

2

1

s

φcmd(autopilot)

Autopilot engaged:

• AP switch in position 2

• Autopilot drives inner-loop 

 controls directly

 

Autopilot disengaged:

• AP switch in position 1

• Pilot uses side arm controller

β

Ka p

~ Ka b

Kaφ

Kaδ

Kap

Kab

r
28

Appendix C

Pilot Written Comments

Abbreviations:

CHR Cooper-Harper rating

HUD head-up display

LFRC load factor rate command

PAL pilot assistance level

RBFD resolver-based flight director

RCAH rate-command–attitude-hold

VMS Visual/Motion Simulator

Pilot A Comments

Experiment Part 1

HUD. Of the HUD options, I prefer the resolverbased flight director (RBFD). The guidance for bank

angle and normal load factor is easy to use and provides

good performance. The reticle (load factor circle) diameter could be larger. If there is a problem with this system,

it occurs when I am trying to close on the g-command

guidance cue while using the rate-command–attitudehold (RCAH) response type. The conventional HUD is a

bit difficult to sort out. Although the director cue can be

tracked, tracking is never really smooth, and one frequently overbanks.

Response type. Pilot performance with the resolverbased flight director was very good, and as stated before,

this flight director seems to be well suited for the cruise

turn task, especially when matched to the load factor rate

command flight control law (LFRC). The LFRC system

enabled the pilot to get out of the loop during the steady

part of the turn after having established the proper bank

angle and load factor. The requirement for an initial forward longitudinal input when starting the roll-out of the

turn was not intuitive and required some pilot adjustment. With the RCAH system, the pilot merely relaxes

aft stick pressure during the roll-out.

Experiment Part 2

PAL 3. Easy, fairly precise, good performance with

some spare time remaining (CHR 2).

PAL 2. Throttle director is too busy; constant correction required. Still good performance in load factor and

bank angle tracking. Control somewhat degraded; little

spare time (CHR 5).

PAL 1. No throttle director; not too shabby. Learning curve remembered for approximate throttle settings.

Real workload involved when bank angle is greater than

45° to 50° on both roll-in and roll-out. Satisfactory performance on bank angle and load factor. Throttle motion

much less than with the throttle director of PAL 2, and

performance with dynamic pressure was almost as good.

A piece of cake during steady segment of turn (CHR 4).

PAL 0. Control of vehicle trajectory will be lost

without flight director (CHR 10).

General comments. For the cruise turns that I performed, task performance clearly deteriorated and my

workload increased with lower PAL settings where less

guidance information was provided. Dynamic pressure

control was particularly erratic because I really didn’t

have an accurate idea of power requirements at lower

PAL’s. Pilot workload during the cruise turns was highest during the roll-in and roll-out, regardless of PAL setting. In PAL’s 2 and 3, the workload during the steady

segment of the turn was very low because the LFRC

flight control system eliminated the longitudinal part of

the task.

Pilot B Comments

Experiment Part 1

Conventional HUD-RCAH flight control law. For

the cruise turn task, this display and flight control law

combination seems well matched and level-1 flying

qualities were achieved with PAL 3. Demands on the

pilot were relatively low during roll-in and roll-out of the

turns, and maintainability of 2g while established in the

Table C1. Cooper-Harper Ratings for PAL 3 (autothrottles on):

[HUD and response type preference: RBFD and LFRC]

HUD Response type Rating

Conventional RCAH 4

Conventional LFRC 3

RBFD RCAH 3

RBFD LFRC 2

Table C2. Pilot A Cooper-Harper Ratings With Preferred HUD

and Response Type

PAL Rating

3 2

2 5

1 4

0 10
29

turn was not difficult. As already noted, discrimination

between the lateral and the longitudinal cues is difficult

when established in the 2g turn (CHR 3).

Conventional HUD-LFRC flight control law. This

combination was obviously not as well matched as the

previous control law; pilot demands were definitely

higher during roll-in and, especially, during roll-out

when the pilot had to push forward to follow the longitudinal cue. After the turn was established at 2g, pilot

demands were low because the pilot could essentially get

out of the loop in the longitudinal axis (CHR 3+).

RBFD-RCAH flight control law. This flight director

and control law combination was the least well matched

of any evaluated. Longitudinal demands on the pilot

were higher during roll-ins, roll-outs, and while established in the turns (CHR 4).

RBFD-LFRC flight control law. This is a good

combination. Although this flight director is not conventional and I had not flown it for some time, I felt very

comfortable with it immediately and my performance

with PAL 3 was definitely level 1. Performance would be

even better if HUD symbology resolution could be

improved to enable better discrimination of the intersection of the lift arrow and the bank angle index. This was

the best flight director and flight control law combination

for the cruise turn task and was therefore chosen for subsequent tasks during this session (CHR 3).

My order of preference for the four combinations is

as follows:

1. RBFD-LFRC

2. Conventional HUD-RCAH

3. Conventional HUD-LFRC

4. RBFD-RCAH

Experiment Part 2

Next the pilot assist level was stepped down from 4

to 0 during the use of the RBFD-LFRC flight control

law.

PAL 4. When completely automatic, I observed

power movements and requirements that the autopilot

used to accomplish the maneuver.

PAL 3. With autothrottles and flight director, performance was the same as in experiment part 1 (CHR 3).

PAL 2. Without the autothrottle, I used the flight

director power-required cue to set power to the desired

level to maintain dynamic pressure to within the specified level. My performance was usually within the

desired criteria, but pilot demands definitely increased,

especially during roll-in and roll-out (CHR 4).

PAL 1. Without the throttle director, I was able to

keep dynamic pressure within the desired range by making power changes in response to deviations and using

guidelines I learned from observing PAL 4, 3, and 2

throttle requirements. My performance for this task with

PAL 1 was probably as good as with PAL 2, but this was

due to my being well up on the learning curve with this

particular task. However, given different initial conditions with different power demands, my performance

would have been much better with PAL 2 (CHR 4).

PAL 0. With the conventional display and LFRC

flight control law, I could not achieve even adequate performance. Altitude could be maintained to within about

1000 ft of target and dynamic pressure within 100 lb/ft2

of target. Load factor varied ±0.5g from the desired factor. Attitude and flight path inconsonance prevented me

from maintaining precise control of the aircraft (CHR 7).

With conventional HUD and RCAH flight control

laws, my performance was only slightly better in terms

of g-control but still well outside of adequate performance criteria (CHR 7).

During this simulation session, my performance of

cruise turns with lower PAL’s appeared to be significantly better than during my previous session. This was

because I had recent practice and was familiar with

power settings and attitudes required to accomplish the

maneuvers.

Table C3. Cooper-Harper Ratings for PAL 3 (autothrottles on):

[HUD and response type preference: RBFD and LFRC]

HUD Response type Rating

Conventional RCAH 3

Conventional LFRC 3+

RBFD RCAH 4

RBFD LFRC 3

Table C4. Pilot B Cooper-Harper Ratings With Preferred HUD

and Response Type

PAL Rating

3 3

2 4

1 4

0 7
30

Pilot C Comments

Experiment Part 1

HUD. The conventional flight director was preferred

over the resolver-based flight director (RBFD) because

errors and error rates in roll and pitch were more easily

perceived. The RBFD was disliked because of symbology clutter which obscured the exact position of the lift

vector tip and required that longitudinal and lateral errors

be somewhat large before they could be detected. The

load factor circle, the lift vector arrowhead, and the bank

angle director form a large fuzzy mass of intersecting

white lines where exact positioning of any single element

cannot be determined.

When a bank angle error exists, the conventional

flight director HUD requires the pilot to apply some

visual extrapolation of the wing line of the flight path

marker (FPM) symbol so that the height of the flight

director diamond above or below the wing level can be

used as a measure of the longitudinal inputs needed. This

process seems to be easy to accomplish, however, and no

noticeable degradation in performance was observed

with this scheme compared with the conceptually better

(in my judgment) RBFD scheme.

When the error in load factor and bank angle are

moderate to small, the conventional flight director gives

superior error and error rate cues, in part, because of having more symbol surfaces and corners to examine for

alignment with each other.

Command response. The NV rate command and NV

hold (LFRC) was preferred to the pitch-rate-command–

attitude-hold (RCAH) system because between roll-in

into the turn and roll-out from the turn, the longitudinal

stick forces could be released once a steady solution was

reached. This permitted more pilot attention to making

bank angle corrections.

Experiment Part 2

PAL settings. The PAL 4 (fully automated) setting

was not very interesting except that a carefully flown

PAL 2 seemed to hold altitude and dynamic pressure

closer to target values. The same result applies to a comparison between the PAL 3 and PAL 2 settings. PAL 2

provided better altitude and dynamic pressure holding.

The cost of the improved performance with PAL 2 is an

increased demand for pilot attention. Any side task loading could result in very different outcomes.

The throttle cues were useful. The throttle gain was

about right for manual control, but its feel was sticky and

abrupt, so steplike changes were made of necessity. In

spite of this, good control of thrust was possible. With

PAL 1, good control was possible just by watching

dynamic pressure values and responding to them.

The PAL 0 setting was possible to fly from the start

of the task up until somewhere during the roll-out on the

new heading. Practice would probably permit that part to

be flown as well. My approach was to try to keep the

flight path marker on the horizon and to adjust bank

angle to keep it there (correcting altitude errors). I set

about 1.98g, let the control system hold it there, and varied thrust around the value that I knew to be about right.

The roll-out was difficult for two reasons. First, my attitude reference on roll-out seemed to go away. Not

enough time was available to understand what was happening here. Secondly, capturing the heading exactly

involved knowing a lead point and a roll-and-push

technique.

General comment. My control strategy was to be

very attentive to small errors and error rates and then try

to quickly and smoothly null them. This required much

attention and would not allow distractions to dilute the

pilot effort applied to the task.

Pilot D Comments

Experiment Part 1

I chose the conventional HUD-LFRC control system. My belief was that the resolver-based flight director

(RBFD) moved the pilot’s focus away from the flight

path of the vehicle because of the compelling nature of

the RBFD load factor circle. The pilot finds himself

striving to fly the lift vector arrow with relation to the

Table C5. Pilot C Cooper-Harper Ratings for PAL 3

(autothrottles on):

[HUD and response type preference: RBFD and LFRC]

HUD Response type Rating

Conventional RCAH 3

Conventional LFRC 2.5

RBFD RCAH 4

RBFD LFRC 4

Table C6. Pilot C Cooper-Harper Ratings With Preferred HUD

and Response Type

PAL Rating

3 2.5

2 3

1 3

0 7
31

load factor circle rather than concentrating on the flight

path. With such narrow flight path deviations acceptable

in hypersonic flight, an omission of gamma as a primary

scan item may prove disastrous. The net effect is to

reduce the overall situational awareness of the pilot. The

conventional HUD with its flight path director permits

the pilot to maintain situational awareness as one flies the

flight path marker to the flight path director.2

The LFRC response type was selected because of the

ease of sustaining the proper load factor and ultimately

the proper altitude during the sustained turn maneuver.

The RCAH response type was inferior in the roll-in and

roll-out of the turn, and the workload was higher during

the sustained turn to maintain altitude control. The LFRC

response type did require forward stick during the rollout, which was not an intuitive input and is considered to

be a minor deficiency of the response type.

Experiment Part 2

The PAL 4 configuration is totally automated, and

therefore, no CHR’s were provided. I had the opportunity

to observe the autothrottle movements (the autothrottles

are backdriven) for eyeball calibration and to observe the

throttle director. PAL 3 was the same configuration

flown to evaluate the most preferred HUD-response type

combination. The desired performance criteria were met,

and I rated the task a CHR of 3. The roll-axis required a

little more effort than was desirable, and the necessity of

pushing forward on the side-arm controller (SAC) during

roll-out was a little annoying. PAL 2 was interesting as

the throttle director appeared to make the task more difficult because it tended to cause me to manipulate the

throttles at a higher gain level with the result of chasing

2Author’s comment: Note that the director symbols for both

HUD symbologies were driven by the same load factor error signal.

When the pilot nulled the flight director error on the conventional

HUD, the pilot was not directly flying the flight path marker to the

director but rather was zeroing the director offset from the flight

path marker by achieving the desired load factor.

the q-bar at times. I found that a better technique for me

was to use the throttle bowtie display in conjunction with

the dynamic pressure display to more smoothly control

the q-bar. Using the throttle director, I gave the task a

CHR of 4 while ignoring the throttle director resulted in

a CHR of 3. This was essentially the PAL 1 configuration. The PAL 0 configuration (no flight director) did not

lend itself to the use of the LFRC response type, so I preferred the RCAH configuration. Because of the compounding problems of centrifugal relief, path and attitude

decoupling, and dynamic pressure sensitivity, just maintaining control was an effort without a flight director.

The CHR of 8 reflects this.

One other comment deserves mention. The motion

cueing of the Langley VMS provided an excessive beta

signal during the cruise turns. The cruise turns are balanced turns, yet the simulator provided a significant sideslip motion cue which was clearly discernible to the

pilot.

Table C7. Pilot D Cooper-Harper Ratings for PAL 3

(autothrottles on):

[HUD and response type preference: RBFD and LFRC]

HUD Response type Rating

Conventional RCAH 4

Conventional LFRC 3

RBFD RCAH 4

RBFD LFRC 3

Table C8. Pilot D Cooper-Harper Ratings With Preferred HUD

and Response Type

PAL Rating

3 3

2 4

1 3

0 8
32

References

1. Walton, James T.: Performance Sensitivity of Hypersonic

Vehicles to Changes in Angle of Attack and Dynamic Pressure. AIAA-89-2463, July 1989.

2. Shaughnessy, John D.; and Raney, David L.: Response of a

Hypersonic Vehicle to Atmospheric Disturbances During

Ascent. NASP TM-1061, NASP JPO, Wright-Patterson AFB,

1989.

3. Berry, Donald T.: National Aero-Space Plane Flying Qualities

Task Definition Study. NASP TM-1039, NASP JPO, WrightPatterson AFB, 1988.

4. Berry, Donald T.: National Aero-Space Plane Flying Qualities

Requirements. NASP TM-1084, NASP JPO, Wright-Patterson

AFB, 1989.

5. McRuer, Duane T.; Ashkenas, Irving L.; and Johnston, Donald

E.: Flying Qualities and Control Issues/Features for Hypersonic Vehicles. NASP CR-1063, NASP JPO, Wright-Patterson

AFB, 1989.

6. Chalk, C. R.: Flying Qualities Criteria Review, Assessment

and Recommendations for NASP. NASP CR-1065, NASP JPO,

Wright-Patterson AFB, 1989.

7. McRuer, Duane T.; and Myers, Thomas T.: Considerations for

the Development of NASP Flying Qualities Specifications.

WL-TR-92-3042, U.S. Air Force, 1991. (Also available from

DTIC as AD B170 015L.)

8. Lallman, Frederick J.; and Raney, David L.: Dynamic

Pressure/Energy Regulator for NASP. NASP TM-1032, NASP

JPO, Wright-Patterson AFB, 1988.

9. Lallman, Frederick J.; and Raney, David L.: Altitude-Velocity

Regulator for NASP. NASP TM-1053, NASP JPO, WrightPatterson AFB, 1989.

10. Lallman, Frederick J.; and Raney, David L.: Energy/Dynamic

Pressure Regulator for NASP. NASP TM-1066, NASP JPO,

Wright-Patterson AFB, 1989.

11. Lallman, Frederick J.; and Raney, David L.: Velocity/Dynamic

Pressure Regulator for NASP. NASA TM-1089, NASP JPO,

Wright-Patterson AFB, 1989.

12. Lallman, Frederick J.; and Raney, David L.: Altitude-Energy

Regulator for NASP. NASP TM-1098, NASP JPO, WrightPatterson AFB, 1990.

13. Raney, David L.; and Lallman, Frederick J.: Control Integration Concept for Hypersonic Cruise-Turn Maneuvers. NASA

TP-3136, 1992.

14. Bonner, E.; Clever, W.; and Dunn, K.: Aerodynamic Preliminary Analysis System II. Part I—Theory. NASA CR-182076,

1991.

15. Shaughnessy, John D.; Pinckney, S. Zane; McMinn, John D.;

Cruz, Christopher I.; and Kelley, Marie-Louise: Hypersonic

Vehicle Simulation Model: Winged-Cone Configuration.

NASA TM-102610, 1990.

16. Hahne, David E.; Luckring, James M.; Covell, Peter F.;

Phillips, W. Pelham; Gatlin, Gregory M.; Shaughnessy, John

D.; and Nguyen, Luat T.: Stability Characteristics of a Conical

Aerospace Plane Concept. SAE Tech. Paper Ser. 892313,

Sept. 1989.

17. Chavez, Frank R.; and Schmidt, David K.: Dynamics of

Hypersonic Flight Vehicles Exhibiting Significant Aeroelastic

and Aeropropulsive Interactions. AIAA-93-3763, Aug. 1993.

18. Chan, Samuel Y.; Cheng, Peter Y.; Pitt, Dale M.; Myers,

Thomas T.; Klyde, David H.; Magdaleno, Raymond E.; and

McRuer, Duane T.: Aeroservoelastic Stabilization Techniques

for Hypersonic Flight Vehicles. NASA CR-187614, 1991.

19. Spain, Charles V.; Zeiler, Thomas A.; Gibbons, Michael D.;

Soistmann, David L.; Pozefsky, Peter; DeJesus, Rafael O.; and

Brannon, Cyprian P.: Aeroelastic Character of a National

Aerospace Plane Demonstrator Concept. AIAA-93-1314,

Apr. 1993.

20. Grantham, William D.: Comparison of Flying Qualities

Derived From In-Flight and Ground-Based Simulators for a

Jet-Transport Airplane for the Approach and Landing Pilot

Tasks. NASA TP-2962, 1989.

21. Martin, D. J., Jr.: A Digital Program for Motion Washout on

Langley’s Six-Degree-of-Freedom Motion Simulator. NASA

CR-145219, 1977.

22. McFarland, Richard E.: A Standard Kinematic Model for

Flight Simulation at NASA-Ames. NASA CR-2497, 1975.

23. Rollins, John D.: Description and Performance of the Langley

Visual Landing Display System. NASA TM-78742, 1978.
33

Table 1. Geometric Characteristics of Vehicle Concept Used in Real-Time Simulation

Wing:

Reference area (includes area projected to fuselage centerline), ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3603

Aspect ratio. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.00

Span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60.0

Leading-edge sweep angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75.97

Trailing-edge sweep angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.0

Mean aerodynamic chord, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80.0

Airfoil section. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Diamond

Airfoil thickness to chord ratio, percent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.0

Incidence angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.0

Dihedral angle, deg. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.0

Wing flap (elevon):

Area (each), ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92.3

Chord (constant), ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.22

Inboard section span location, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15.0

Outboard section span location, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27.78

Vertical tail, body centerline:

Exposed area, ft2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 645.7

Theoretical area, ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1248.8

Span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32.48

Leading-edge sweep angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70.0

Trailing-edge sweep angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38.13

Airfoil section. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Diamond

Airfoil thickness to chord ratio, percent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.0

Rudder:

Area, ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 161.4

Span, ft2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22.8

Chord to vertical tail chord ratio, percent. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25.0

Axisymmetric fuselage:

Theoretical length, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200.0

Cone half angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.0

Cylinder radius (maximum), ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.87

Cylinder length, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.88

Boattail half angle, deg. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.0

Boattail length, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40.0

Moment reference center, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124.01
34

Table 2. Performance Limits of Langley Visual/Motion Simulator

[See ref. 20]

Degree of freedom Position, deg Velocity, deg/sec Acceleration, deg/sec2

Pitch . . . . . . . . . . . . . +30 to −20 ±15 ±50

Roll . . . . . . . . . . . . . ±22 ±15 ±50

Yaw . . . . . . . . . . . . . ±32 ±15 ±50

Position, ft Velocity, ft/sec Acceleration, g units

Vertical . . . . . . . . . . +2.5 to −3.25 ±2 ±0.6

Lateral . . . . . . . . . . . ±4.0 ±2 ±0.6

Longitudinal. . . . . . . +4.1 to −4.0 ±2 ±0.6

Table 3. Feel System Settings of McFadden Control Loader Used for Side Stick

Parameter Pitch axis Roll axis

Breakout force, lb . . . . . . . . . ±1.0 0.75 left, 1.0 right

Displacement limits, deg . . . . 18 fwd, 20 aft ±20

Maximum force, lb . . . . . . . . 12 fwd, 13 aft 5.0 left, 6.5 right

Velocity limit, in/sec . . . . . . . 35 35

Frequency, rad/sec . . . . . . . . . 23 16

Damping . . . . . . . . . . . . . . . . 0.85 1.4

Table 4. Task Completion Criteria Used in Hypersonic Cruise Turn Maneuver

Parameter Completion criteria

Heading Maintain within ±2° of target heading for 5 sec

Bank angle Maintain magnitude of φ < 3° for 3 sec

Altitude Maintain within ±80 ft of target altitude for 5 sec

Dynamic pressure Maintain within ±20 lb/ft2 of target value for 5 sec
35

Table 5. Definition of Pilot Assistance Level (PAL) Designations

PAL designation Flight director Throttle director Autothrottle Autopilot Description

0 OFF OFF OFF OFF Manual control, unassisted

1 ON OFF OFF OFF Manual control with flight

director

 2 ON ON OFF OFF Manual control with flight

and throttle directors

 3 ON ON ON OFF Manual stick with autothrottle and both directors

 4 ON ON ON ON Fully automated control:

autopilot and autothrottle

Table 6. Designations for Response Type and HUD Combinations

HUD symbology

Control response type Conventional RBFD

RCAH RCAH-1 RCAH-2

LFRC LFRC-1 LFRC-2

Table 7. Pilot Preferences for Response Type and HUD Combination

Response type HUD symbology Preferred

Pilot RCAH LFRC Nominal RBFD combination

A √ √ LFRC-2

B √ √ LFRC-2

C √ √ LFRC-1

D √ √ LFRC-1
36

Figure 1. Hypersonic vehicle concept used in real-time simulation. All dimensions are in feet.

Figure 2. Variation of vehicle inertial characteristics and center of gravity with weight.

Figure 3. Interior layout of simulator cockpit.

Elevon

Moment reference center Rudder

60

124

200

2

4

6

8

10

9

10

11

12

13

14

15

1.0 1.5 2.0 2.5 3.0

Ix

Iy

Iz

xc.g.

0

Weight, lb

× 106

× 105

xc.g., ft

Moment of

inertia,

slug-ft2

EADI SPI EADI

HUD and CGI

CDU

SPI

HSI HSI

Left front monitor Right front monitor

Left side monitor

HUD and CGI

CGI

Left station Right station

Side arm

controller

Armrest Throttle 

levers

Electronic attitude display indicator

Horizontal situation indicator

Computer-generated imagery

. . . . .

. . . . .

. . . . .

EADI

HSI

CGI

Surface position indicator

Head-up display

Cockpit display unit

. . . . .

. . . . .

. . . . .

SPI

HUD

CDU
37

Figure 4. Symbology for baseline head-up display used in real-time simulation.

Figure 5. Diagram of head-down electronic attitude display indicator (EADI).

768 85000

3.7 α

M

G

q

7.85

2000

0.84

Altitude, ft

Heading indicator

Artificial horizon line

Flight path marker

Pitch ladder

Throttle level indicator

Equivalent

airspeed, knots

Pitch attitude indicator

5

5

10

5

5

10

080 090 100

10 10

Bank angle indicator

15

10

5

5

5

5

10

15

258

Flight path marker

Pitch 

ladder

Lateral acceleration indicator

15000

Altitude, ft

Artificial horizon line

Heading

M .54

A 7.6

Roll index

30 30

2.5

347

Equivalent airspeed, knots

2.0

1.0

.5

1.2 G

Mach

Alpha

Normal

accel

tape

Sink

rate

tape

Load factor
38

Figure 6. Diagram of head-down horizontal situation indicator (HSI) display.

Figure 7. Diagram of head-down surface position indicator (SPI) display.

258

190/12

COURSE

270

HDG TRU

N

3

6

12

15

21

24

30

S

33

W

E

Wind vector

indicator

Fuel consumed, lb

6, 479

Fuel flow

rate, lb/sec

Fuel remaining, lb

103,521

Weight, lb 243,521

 Nominal: Green

 Rate-limited: Yellow

 Position-limited: Red

Color code for control surface deflections

Fill line, moves downward as 

fuel is depleted

123

–4.2° +8.3°

+3.1°
39

Figure 8. Time history of two typical cruise turn maneuvers executed during preliminary simulation sessions at

Mach 7.86 and 85 000 ft.

Pilot A

Pilot B

Time, sec

Pilot A

Pilot B

Pilot A

Pilot B

Pilot A

Pilot B

Angle of attack, deg

Pilot A

Pilot B

Altitude, ft 

Pilot A

Pilot B

Fuel flow rate, lb/sec

Heading, deg Dynamic pressure, lb/ft2 Bank angle, deg

Time, sec

Time, sec

Time, sec

Time, sec

Time, sec
40

(a) Pilot A.

(b) Pilot B.

Figure 9. Variation in pitch attitude and flight path angle for typical cruise turn maneuvers executed during preliminary

simulation sessions at Mach 7.86 and 85 000 ft.

Pitch attitude variation

Flight path angle

Time, sec

Angle, deg

Pitch attitude variation

Flight path angle

Time, sec

Angle, deg
41

Figure 10. Symbology for conventional flight director.

Figure 11. Symbology for resolver-based flight director (RBFD).

Flight director

Pitch attitude indicator

768 85000

3.8 α

M

G

q

7.85

2000

1.20

5

5

10

15

5

5

10

15

050 090 130

Altitude, ft

Actual heading indicator

Artificial horizon line

Flight path marker

Pitch ladder

Equivalent

airspeed, knots

Desired heading indicator

Throttle director

85000

5

5

10

15

5

5

10

15

050 090 130

768

3.8 α

M

G

q

7.85

2000

1.20

Bank angle director

Lift vector arrow

Nominal bank indicator

Load factor circle

Artificial horizon line

Altitude, ft

Actual heading indicator

Flight path marker

Pitch ladder

Throttle director

Equivalent

airspeed, knots

Desired heading indicator
42

Figure 12. Flowchart showing elements of real-time simulation setup.

Figure 13. Cooper-Harper flying qualities rating scale.

Task inputs

from CDU

Outer-loop control elements

Inner-loop flight controls

 Inner-loop feedbacks 

Autopilot 

Autothrottle 

Hypersonic vehicle model

Aero model

Propulsion model

Inertial model

Atmospheric model

Kinematic equations

Flight dynamic equations

Auxiliary relations

Yaw 

control

Roll 

control

Flight 

director

Pilot

Stick

Throttle

levers Outer-loop feedbacks 

Pitch 

control

( LFRC 

or

RCAH)

Autopilot

Autothrottle

A

Simulation 

outputs 

ON

OFF

OFF

ON

OFF

ON

ON

OFF

Throttle

director

Aircraft 

characteristics 

Demands on the pilot in 

selected task or required operation*

Pilot

rating

Excellent 

Highly desirable

Good 

Negligible deficiencies 

Fair - Some mildly

unpleasant deficiencies

Minor but annoying 

deficiencies 

Moderately objectionable 

deficiencies 

Very objectionable but 

tolerable deficiencies 

Major deficiencies

Major deficiencies

Major deficiencies

Major deficiencies 

Pilot compensation not a factor 

for desired performance

Pilot compensation not a factor 

for desired performance

Minimal pilot compensation 

required for desired perfomance 

Desired performance requires 

moderate pilot compensation

Adequate performance requires 

considerable pilot compensation 

Adequate performance requires 

extensive pilot compensation 

Adequate performance not attainable 

with maximum tolerable pilot 

compensation controllability not in

question 

Considerable pilot compensation is 

required for control 

Intense pilot compensation is 

required to retain control 

Control will be lost during some 

portion of required operation 

1

2

3

4

5

6

7

10

8

9

Deficiencies 

warrant 

improvement 

Is it 

satisfactory 

without improvement?

Deficiencies 

require 

improvement

Is adequate

performance 

attainable with 

a tolerable 

pilot workload?

Improvement 

mandatory 

Is it 

controllable?

Pilot

decisions

Yes

No

No

No

Yes

Yes

Adequacy for selected task 

or required operation*

*Definition of required operation involves designation of flight phase 

and/or subphases with accompanying conditions
43

Figure 14. Itinerary for typical real-time simulation session.

Real Time Session -- Itinerary Hypersonic Simulation

 Pilot: Date:

Session itinerary for up & away maneuver tasks:

Part I: Execute cruise-turn comparing 4 Response Type/ HUD combinations using

PAL 3; select preferred combination.

 a) Command Response Types: RCAH, LFRC

 b) HUD Types: HUD 0a (conventional), HUD 2 (RBFD)

Part II: Execute cruise-turn using successively lower PAL settings with preferred

response type and HUD combination; note perfomance degradation with PAL

reduction.

 c) PAL Settings:

PAL Flight Director Throttle Director Autothrottle Autopilot

4X X X X

3 X X X --

2 X X -- --

1 X -- -- --

0 -- -- -- --

Cruise-Turn Task Defintion: 30 deg. heading change; 2-g turn from 090 to 120

 (Initiate from level flight at Mach= 7.90, dynamic pressure= 2000 psf, h= 85,040 ft)

Hold dynamic pressure and altitude

Bank to initiate maneuver

Establish steady 2g turn

Attain required heading change

Roll out on desired heading

Flight Director Tracking Tolerances:

 Desired Adequate

Normal Load Factor ± 0.10 g ± 0.20 g

Bank Angle ± 5.0 deg ± 10.0 deg

Target Parameter Acquisition/ Regulation Tolerances:

Desired Adequate

Dynamic Pressure ± 20.0 psf ± 30.0 psf

Altitude ± 200 ft ± 300 ft

Heading ± 0.5 deg ± 1.0 deg
44

Figure 15. Cooper-Harper ratings of response type and HUD combination with PAL 3.

Figure 16. Stick deflection (rms) versus response type and HUD combination with PAL 3.

1.5

2.0

2.5

3.0

3.5

4.0

4.5

Pilot A

Pilot B

Pilot C

Pilot D

Level 1

Level 2

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

CHR

Max.

Min.

Avg.

.1

.2

.3

.4

.5

.6

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

Stick deflection, in. rms
45

Figure 17. Time histories of typical maneuvers conducted by pilot A using each of four response type and HUD

combinations.Dynamic pressure, lb/ft2

RCAH-1

LFRC-1

RCAH-2

LFRC-2

Desired performance bounds: 2000 + 20 lb/ft2

Adequate performance bounds: 2000 + 30 lb/ft2

Time, sec

RCAH-1

LFRC-1

RCAH-2

LFRC-2

Altitude, ft

Desired performance bounds: 85 040 + 200 ft

Adequate performance bounds: 85 040 + 300 ft

Time, sec
46

Figure 17. Continued.

RCAH-1

LFRC-1

RCAH-2

LFRC-2

Time, sec

Normal load factor, g

Time, sec

Angle of attack, deg

RCAH-1

LFRC-1

RCAH-2

LFRC-2
47

Figure 17. Concluded.

Time, sec

Bank angle, deg

RCAH-1

LFRC-1

RCAH-2

LFRC-2

Time, sec

Heading, deg

RCAH-1

LFRC-1

RCAH-2

LFRC-2
48

Figure 18. Normal load factor rms tracking error versus response type and HUD combination with PAL 3.

Figure 19. Bank angle rms tracking error versus response type and HUD combination with PAL 3.

0

.05

.10

.15

Load factor rms tracking error, 

g

Max.

Min.

Desired: error < 0.1 Avg. g

Adequate: error < 0.2g

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

Bank angle rms tracking error, deg

Max.

Min.

Desired: error < 5° Avg.

Adequate: error < 10°

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination
49

Figure 20. Peak absolute dynamic pressure error versus response type and HUD combination with PAL 3.

Figure 21. Mean absolute dynamic pressure error versus response type and HUD combination with PAL 3.

0

5

10

15

Peak dynamic pressure error, lb/ft2

Max.

Min.

Desired: error < 20 lb/ft Avg.

2

Adequate: error < 30 lb/ft2

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

0

1

2

3

4

Mean dynamic pressure error, lb/ft2

Max.

Min.

Avg.

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

Desired: error < 20 lb/ft2

Adequate: error < 30 lb/ft2
50

Figure 22. Peak absolute altitude error versus response type and HUD combination with PAL 3.

Figure 23. Mean absolute altitude error versus response type and HUD combination with PAL 3.

0

50

100

150

200

Peak altitude error, ft

Max.

Min.

Desired: error < 200 ft Avg.

Adequate: error < 300 ft

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

0

20

40

60

80

Mean altitude error, ft

Max.

Min.

Avg.

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

Desired: error < 200 ft

Adequate: error < 300 ft
51

Figure 24. Fuel consumption versus response type and HUD combination with PAL 3.

Figure 25. Cooper-Harper ratings with pilot assistance levels.

6050

6100

6150

6200

6250

6300

6350

6400

Fuel consumption, lb

6791 lb

Max.

Min.

Avg.

RCAH-1 LFRC-1 LFRC-2 RCAH-2

Response type-HUD combination

0

2

4

6

8

10

12

CHR

PAL

0123

Pilot A

Pilot B

Pilot C

Pilot D

Level 3

Level 2

Level 1
52

Figure 26. Stick deflection (rms) versus pilot assistance level.

Figure 27. Time histories of longitudinal stick deflection from two maneuvers conducted by pilot B using the LFRC

response type with PAL 0 and PAL 3.

–.1

0

.1

.2

.3

.4

.5

.6

01234

Stick deflection, in. rms

PAL

Max.

Min.

0.72 in. Avg.

PAL 0

PAL 3

Longitudinal stick deflection, in.

PAL 0

PAL 3

Time, sec
53

Figure 28. Peak absolute dynamic pressure error versus pilot assistance level.

Figure 29. Peak absolute altitude error versus pilot assistance level.

0

20

40

60

80

100

120

01234

Peak dynamic pressure error, lb/ft2

PAL

 Desired performance 

 Adequate performance 

223 lb/ft2

Max.

Min.

Desired: error < 20 lb/ft Avg.

2

Adequate: error < 30 lb/ft2

0

50

100

150

200

01234

Peak altitude error, ft

PAL

Max.

Min.

Desired: error < 200 ft Avg.

Adequate: error < 300 ft

Max. = 3860 ft

Min. = 571 ft

Avg. = 1740 ft
54

Figure 30. Mean absolute dynamic pressure error versus pilot assistance level.

Figure 31. Mean absolute altitude error versus pilot assistance level.

0

10

20

30

40

50

01234

Mean dynamic pressure error, lb/ft2

PAL

 Desired performance 

 Adequate performance 

110 lb/ft2

Max.

Min.

Desired: error < 20 lb/ft Avg.

2

Adequate: error < 30 lb/ft2

0

20

40

60

80

Mean altitude error, ft

01234

PAL

Max.

Min.

Desired: error < 200 ft Avg.

Adequate: error < 300 ft

Max. = 2015 ft

Min. = 198 ft

Avg. = 802 ft
55

Figure 32. Time histories of typical maneuvers conducted by each of four pilots in PAL 0 using their preferred response

type and HUD combination.

Time, sec

Pilot A

Pilot B

Pilot C

Pilot D

Dynamic pressure, lb/ft2

Desired performance bounds: 2000 + 20 lb/ft2

Adequate performance bounds: 2000 + 30 lb/ft2

 Desired

 Adequate

 Desired

Altitude, ft

Time, sec

Pilot A

Pilot B

Pilot C

Pilot D

 Desired

 Adequate

 Desired

 Adequate

 Adequate

Desired performance bounds: 85 040 + 200 ft

Adequate performance bounds: 85 040 + 300 ft

Desired performance bounds: 85 040
56

Figure 32. Continued.

Time, sec

Normal load factor, g

Pilot A

Pilot B

Pilot C

Pilot D

Time, sec

Angle of attack, deg

Pilot A

Pilot B

Pilot C

Pilot D
57

Figure 32. Continued.

Time, sec

Bank angle, deg

Pilot A

Pilot B

Pilot C

Pilot D

Time, sec

Heading, deg

Pilot A

Pilot B

Pilot C

Pilot D
58

Figure 32. Concluded.

Time, sec

Pilot A

Pilot B

Pilot C

Pilot D

Fuel flow rate, lb/sec
59

Figure 33. Time histories from typical maneuvers conducted by pilot D with all pilot assistance levels.

hdg

Time, sec

Dynamic pressure, lb/ft2

Desired performance bounds: 2000 + 20 lb/ft2

Adequate performance bounds: 2000 + 30 lb/ft2

 Desired

 Adequate

 Adequate

 Desired

PAL 4

PAL 3

PAL 2

PAL 1

* PAL 0

*

*

*

Altitude, ft

Time, sec

 Desired

 Adequate

 Desired

 Adequate

PAL 4

PAL 3

PAL 2

PAL 1

* PAL 0

*

*

*

Desired performance bounds: 85 040 + 200 ft

Adequate performance bounds: 85 040 + 300 ft

Desired performance bounds: 85 040
60

Figure 33. Continued.

Time, sec

Normal load factor, g

PAL 4

PAL 3

PAL 2

PAL 1

PAL 0

*

*

*

*

Time, sec

Angle of attack, deg

PAL 4

PAL 3

PAL 2

PAL 1

PAL 0

*

*

*

*
61

Figure 33. Continued.

Time, sec

Bank angle, deg

PAL 4

PAL 3

PAL 2

PAL 1

* PAL 0

*

*

*

Time, sec

Heading, deg

PAL 4

PAL 3

PAL 2

PAL 1

* PAL 0

*

*

*
62

Figure 33. Concluded.

Time, sec

Fuel flow rate, lb/sec

PAL 4

PAL 3

PAL 2

PAL 1

PAL 0 *

*

*

*
63

Figure 34. Fuel consumption versus pilot assistance level.

6000

6200

6400

6600

6800

7000

7200

7400

7600

01234

Fuel consumption, lb

PAL

9230 lb 

Max.

Min.

Avg.
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REPORT DOCUMENTATION PAGE

October 1995 Technical Paper

Investigation of Piloting Aids for Manual Control of Hypersonic Maneuvers WU 505-70-64-01

David L. Raney, Michael R. Phillips, and Lee H. Person, Jr.

L-17446

NASA TP-3525

An investigation of piloting aids designed to provide precise maneuver control for an air-breathing hypersonic

vehicle is described. Stringent constraints and nonintuitive high-speed flight effects associated with maneuvering

in the hypersonic regime raise the question of whether manual control of such a vehicle should even be considered.

The objectives of this research were to determine the extent of manual control that is desirable for a vehicle maneuvering in this regime and to identify the form of aids that must be supplied to the pilot to make such control feasible. A piloted real-time motion-based simulation of a hypersonic vehicle concept was used for this study, and the

investigation focused on a single representative cruise turn maneuver. Piloting aids, which consisted of an autothrottle, throttle director, autopilot, flight director, and two head-up display configurations, were developed and

evaluated. Two longitudinal control response types consisting of a rate-command–attitude-hold system and a loadfactor-rate–load-factor-hold system were also compared. The complete set of piloting aids, which consisted of the

autothrottle, throttle director, and flight director, improved the average Cooper-Harper flying qualities ratings from

8 to 2.6, even though identical inner-loop stability and control augmentation was provided in all cases. The flight

director was determined to be the most critical of these aids, and the cruise turn maneuver was unachievable to

adequate performance specifications in the absence of this flight director.

Hypersonic maneuvers; Automatic flight control; Hypersonic vehicles; Hypersonic

flight; Flight dynamics; Hypersonic flying qualities

64

A04

NASA Langley Research Center

Hampton, VA 23681-0001

National Aeronautics and Space Administration

Washington, DC 20546-0001

Unclassified–Unlimited

Subject Category 08

Availability: NASA CASI (301) 621-0390

Unclassified Unclassified Unclassified