~ 

1 

I 

" 

;NASA Contralctor Report 170407 

NASA-CR-170407 

19840012497 

~Spalce ;Shuttle Flying Qualities 

:Crit_:!riBl Assessment - Phase III 

T. T. tvlyers, D. E. Johnston, and D. T. McRuer 

Contract NAS4·3005 

February 19€!4 

NM;/\ 

National Aeronautics and 

Space Administration 

111111111111111111111111111111111111111111111 

NF02560 
NASA Contractor Fleport 170407 

Spclce Shuttle Flying Qualities 

Cri1teria Assessment - Phase III 

T. T. Myers, D. E .• johnston, and D. T. McRuer 

Systems Technology, Inc., 13766 Hawthorne Boulevard, Hawthorne, California 90250 

Prepared for 

Ames Research Center 

Dryden Flight Research Facility 

Edwards, California 

under Contract NAS4-3005 

1984 

NI\S/\ 

National jlleronautics and 

Space Aeronautics 

Ames Reslearch Center 

Dryden Flight Research Facility 

Edwards, California 93523 
This Page Intentionally Left Blank 
I .. 

II.. 

Ill. 

IV., 

V., 

TABLE OF CONTENTS 

INTRODLJCrf ION •••••••••••••••••••••••••••••••••••••••• G •••••••• 1 

PILOT /VEHICLE IDENTIFICATION FROM FLIGHT DATA •••••••••••••••• 5 

A. Overview of Flight Data Analysis ..•.•...•...••••••••.••.. 5 

B. Identification of the Effective Augmented 

Vehicle from Fight Data.................................. 9 

C. Pilot Technique Identification •..•.••••••••••••••••••.•.• 28 

D. Summary of Flight Data Problems and 

Recommendations for OFQ •••.•.••••.•••••••.......•.•..•••. 46 

CREW QUALITATIVE ASSESSMENTS ••••••••••••••••••••••••••••••••• 65 

A. 81'8-5 and -6 Crew Commentary ••••••••••••••••••••••••••••• 66 

B. Revised Crew Interview/Debriefing Questionnaire •••••••••• 69 

HAl'll) CONTROLLER EXPERIMENT ••••••••••••••••••••••••••••••••••• 71 

A. Key Concepts for Assessing Sensitivity 

to Task Variables........................................ 72 

B. Experimental Plan ......•.•.•••...•....•.•..•.••.••••.•..• 

c. Results ...•.••.•..•...•.........•••••••.. 0 ••••••••••••••• 

D. Conclusions .................•...•.••..•...•.. 0 • ••••••••••• 

RESULTS, CONCLUSIONS, AND RECOMMENDATIONS •••••••••••••••••••• 

A. Ma j 0 r Re s u 1 t s •••••••••••••••••••••••••••••••••••••••••••• 

B. Cancius ions .....•...•.............•.•.•.. 0 ••••••••••••••• 

c. Recommendations for OFQ ••••••••••••••••••••••.••••••••••• 

77 

79 

81 

83 

83 

84 

85 

RE~'E:REN(:~~S ••••••••••••••••••••••••••••••••••••••••••••••••••••••• •• R-l 

APPENDIX A. DERIVATION OF NIPIP ESTIMATION EQUATION 

FOR THI<~ SHALLOW GLIDE AND FLARE MODEL................. A-I 

APPENDIX B. SPACE SHUTTLE FLYING QUALITIES QUESTIONNAIRE .......... B-1 

TR-l197-1 iii 
LIST OF FIGURES 

1. Shuttle Approach and Landing (Nominal Characteristics)........ 6 

2. Attitude Control Time Histories 8T8-4......................... 8 

3. 8uperaugmentation Concept..................................... 11 

4. Shuttle Airframe Longitudinal Stability, FlexibleApproach and Landing (Adapted from Ref. 6) •••••••••••••••••••• 14 

5. Schedule for Pitch Rate Loop Gain, Kq (GDCLCOHP) .............. 16 

6. Effective Pitch Gain Scheduling of Shuttle •••••••••••••••••••• 17 

7. FREDA Output for q/oRHC ST3-4 Preflare 

Through Tone hdown .•..•..••• e • " •••• ~ ••••••••••••••• 0 •••••••• 0 • • 20 . 

8. E'RElJA Output for q/oRHC' STS-4 Shallow 

Glide and Final Flare ••••••••••••••••••••••••••••••••••••••••• 22 

9. Definitiol1 of Cost Function in MFP Program •••••••••••••••••••• 23 

10. Comparison of Flight-Derived Effective q/oRHC 

witll l/Tq Fixed and Free •••••••••••••••••••••••••••.•••••••••• 24 

ll. Asymptotic Comparison of Iq/ORHCI with 

Three Different Numet"ator Constraints......................... 27 

12. Comparison of NIPIP and FREDA Results for 

Shuttle Effective Vehicle Identification •••••••••••••••••••••• 29 

11. Elements of Pre flare Oscillation Analysis ••••••••••••••••••••• 31 

14. Iderllized Trajectory for the Shallow Glide 

and Flare Trajectory Model •••••••••••••••••••••••••••••••••••• 33 

15. Effect of Pilot's Flare Time Constant on Final Flare •••••••••• 35 

16. Shallow Glide and Final Flare Pilot Model..................... 36 

17. Idealized Altitude/Sinkrate Phase Plane Trajectory 

for the Shallow Glide and Flare Pilot Model ••••••••••••••••••• 38 

18. Altitude-Sinkrate Phase Plane Analysis of STS-4 Shallow 

Glide and Final Flare Complementary Filtered ~ •••••••••••••••• 39 

19. Variations of Touchdown Variables in the Pilot's 

Tf-Yo Control Plane, STS-4 •••••••••••••••••••••••••••••••••••• 42 

TR-1l97-1 iv 
LIST OF FIGURES (Concluded) 

20. Y p Prom NIl:>IP Iaterpreted in w' Domain ••.••••••••••••••••••••• 

e 

21. Comparison of Altitude Time Histories from Radar 

and IHU Channels i.n DFRF ST3-4 MMLE Files ••••••••••••••••••••• 

22. Interpolated Radar Alt itude, Shallow Glide 

and Final Flare............................................... 50 

23. Comparison of Sinkrate Signals ••••.•..•.•.••..•.••.•...••..•.• 52 

24. Sinkrate from Complementary Filter ••••••••.••.••.••.•••••..•... 53 

25. Altitude-Sinkrate Trajectory, STS-4 Shallow Glide 

and Final Flare, ~ from Reference 18 •••••••••••••••••••••••••• 55 

26. Frequency Response Comparison of e/oRHC from 

Attitude Processor and ACIP •.•••••••••••••••••••.••.•••••..••. 56 

27. NASA .JSC i'1aster Product Data Base for Shuttle................. 60 

28. Typical Page from the STS-6 Computer Compatible 

Tape (CCT) Report for the NASA JSC MPD3 •.•••..••••••••••.....• 61 

29. Generation and Use of the OFQ Archive Files •••••••••••••••••.. 63 

30. Shuttle Orbiter RHC Feel/Command Gradient 

Characteristics in Landing ••.•••••••••••••••.•••••••.•••••.••• 72 

11. The "Critical-Instability" Task (First Order) ••.•••••••••••••• 74 

32. Structural-Isomorphic Hodel of Pilot-Aircraft System •••••.•••• 75 

33. Critical Instability vs. Type of Control Stick................ 76 

34. R.HC Configuration ... 0......................................... 78 

35. Manipulator Experiment Set-Up ••••••••••••••••••••••••••• · •• : •• o. 78 

36. Pitch Axis CCT Raw Scores..................................... 80 

.37. Roll Axis GeT 1\aw Scores .....••••• G ••• 0 ••••••••••••• t ••••••••• 80 

38. Effective Pilot Latency; Shuttle vs. DFRF 

Simulation Controller ••••••••••••••••••••••••••••••••• ~ ••••••• 132 

39. Comparison of Experimental Results with Free and 

Sti ff ~'1anipulator Data of Reference 13........................ 82 

TR-1l97-1 v 
LIST OF TABLES 

1. Comparison of Parameter Extracted from 8TS-4 Flight 

Data with the Superaugmentation Model •••••••••••••••••••••••••• 25 

2. Contribution of FCS Elements to Effective Time Delay ••••••••••• 26 

'3. Shallow Glide and Flare Parameters •••••.•••••..••••.••••••••••• 41 

4. Summary of Shuttle Flight Data Desired and 

Sources for OFQ •••••••••••••••••••••••••••••••••••••••• & • 0 ••••• 58 

TR-1l97-1 vi 
A x, A y' A 

7-

c 

CD 

CD 

0 

TR-1197-1 

NOMENCLATURE 

Acceleration components along the body x, y, z axes 

Reference chord length 

2D Drag coefficient; ----

pU2S 

Zero lift drag coefficient 

21 1ift coefficient; ----

pU2S 

1ift coefficient, a 

a(cic/2U) 

'a (qc!2U) 

ae L 

a8 

o 

2M Pitching moment coefficient; -----

pU2Sc 

vii 
3(ac/2U) 

CMo 

d Beam deviation 

D Aerodynamic drag 

Specific kinetic energy 

Nyquist frequency, Hz 

FD Flight director signal 

g Acceleration due to gravity (32.2 ft/sec 2) 

Washout equalization 

h Path deviation; altitude 

Desired path 

Path error 

Flare initiation altitude 

Altitude at start of shallow glide 

Depth of flare asymptote 

Altitude from IMU 

Radar alt itude 

" 

. 

hTD Touchdown sinkrate 

Moment of inertia in pitch 

Gain 

TR-1197-1 viii 
K "q 

v· r,v 

L 

m 

H 

1'1 u 

n 

TR-1l97-1 

Controller gain 

Averaged value of -dV/dt 

Aerodynamic lift 

Distance between pilot and e.g., positive for pilot 

forward 

Aircraft mass 

Aerodynamic pitching moment; Hach number 

(Single degree of freedom) pitch damping; Pitching 

acceleration per unit pitching velocity; 

Pitching acceleration per unit forward velocity; 

Weathercock stability; Pitching acceleration per unit 

vertical velocity (angle of attack); 

pSUc 

-- CM 21y a 

Pitching acceleration per unit angle of attack rate 

pSUc 2 

4Iy CMa 

Pitching acceleration per unit control surface 

deflection 

Load factor 

Load factor in z direction 

ix 
P, Q, R 

q 

qe 

qss 

q 

R 

Ro 

s 

s 

, 

tTD 

T 

Tf 

Thl' Th2' Th3 

Tq 

Tr 

TRUN 

Te l' Te 2 

TR-1l97-1 

Altitude-to-elevator numerator 

Pitch attitude-to-elevator numerator 

Angular velocity components about the x, y, z axes 

Pitching velocity (pertubation) 

Pitching velocity error 

Steady-state pitching velocity 

Dynamic pressure 

Radius, (Fig. 1) 

Equilibrium yaw rate 

Laplace Operator 

Reference planform area 

Touchdown time measured from flare initiation 

Time constant; sampling period 

Flare time constant 

Time constants, altitude numerator 

Lead time constant in augmentaion system 

Rise time 

Length of a time response for parameter 

identification 

Time constants, real short period roots 

Time constants, pitch attitude numerator 

Lead time constant in short-period e/8 transfer function for conventional airplane dynamics. Lag time 

constant between flight path angle, y, and pitch 

attitude, e. 

x 
u 

V 

w 

w 

TR-1l97-1 

Speed (x stability axis) 

Speed in reference condition 

Speed 

Speed at flare initiation 

Velocity at start of shallow glide 

Equilibrium speed 

Touchdown speed 

Vertical velocity 

Vertical gust velocity 

Gross weight; Aircraft weight 

Flare distance 

Forward acceleration per unit speed change 

p SU (-CD - CD ) 

m u 

Forward acceleration per unit vertical velocity 

= 

pSU 

- (CL - CD ) 2m a 

Controlled element (augmented Shuttle) describing 

function 

Pilot describing function for pilot control action on 

path deviation 

Pilot describing function for pilot control action on 

attitude 

Vertical acceleration per unit forward velocity 

Heave damping; vertical acceleration per unit 

vertical velocity (angle of attack); 

Z a 

xi 

pSU ( ) 2m -CLa - CD 
Y 

Yo 

o 

°RHC 

TR-1197-1 

VertLcal acceleration per unit control surface 

de flection 

Angle of attack 

Aerodynamic angle of attack, (w - wg)/Uo 

Inertial angle of attack, w/Uo 

Sideslip angle, v/Uo 

Flight path angle 

Selected flight path angle, Fig. 22 

Generic control (elevator, horizontal tail, etc.) 

surface deflection 

Aileron deflection 

Body flap deflection 

Body flap controller 

Elevator deflection 

Pilot command input 

Rudder pedal deflection 

Rotational hand controller deflection 

Rudder deflection 

Speed brake 

Speed brake controller 

Throttle deflection 

Distance traveled in flare 

Distance between pilot and ICR, positive for pilot 

forward 

Visual angle error 

xii 
I;; 

I;; sp 

8 

. 

8 

8 e 

8 e 

p 

(j 

T 

T 11 

Tp 

q) , <P 

q)rnm(w) 

q) mn ( j.u) 

<PH 

If " 8 , q) 

W 

we 

wea 

wmin 

wn 

wSP 

TR-1197-1 

Damping ratio 

Short period damping ratio 

Pitch attitude (perturbation) 

Pitch attitude rate 

Attitude Command 

Attitude Error 

Atmospheric Density; coherence function (Eq. 7) 

Real component of complex variable, s 

Time delay 

Delay margin, <PM/wc 

Pilot's time delay 

Bank Angle 

Power spectral density functions 

Cross-spectral density functions 

Phase Hargin 

Euler angles to aircraft body axes (heading, pitch, 

roll sequence) 

Imaginary component of complex variable, s 

Crossover Frequency 

Crossover frequency of amplitude ratio asymptote 

Theoretical lower bound on an identified frequency 

response, 2IT/TRUN 

Natural frequency 

Short period frequency 

xiii 
ACIP 

ADI 

ADS 

AFFTC 

AGL 

ALT 

AMI 

APU 

AR 

ARC 

AVVI 

BET 

BFCS 

CCIT 

CHPR 

CSS 

CTT 

DAP 

DFA 

DFI 

DFRF 

EPR 

FAR 

FBW' 

TR-1l97-1 

ACRONYMS AND ABBREVIATIONS 

Aerodynamic Coefficient Identification Package 

Attitude director instrument 

Air data system 

Air Force Flight Test Center 

Above ground level 

Approach and Landing Test 

Shuttle display (a, Ax, M, V) 

Auxiliary power unit 

Aspect ratio 

fulles Research Center 

Shuttle display (h, h, h) 

Best estimated trajectory 

Backup flight control system 

Cross coupled instability task 

Cooper Harper Pilot Rating 

Control stick steering 

Critical task tester 

Digital autopilot 

Describing function analysis 

Developmental flight instrumentation 

Dryden Flight Research Facility 

Eye-point-of-regard 

Federal Air Regulations 

Fly-by-wire 

xiv 
FCS 

FD 

FREDA 

FSAA 

FSL 

FWD 

GPC 

HAC 

HSI 

HUD 

Hz 

ICR 

IMU 

JSC 

KEAS 

LAHOS 

LOES 

LRC 

MMLE 

MPDB 

NIPIP 

NLR 

OEX 

OFT 

OFQ 

01 

TR-1l9l-1 

Flight control system 

Flight director 

Frequency domain analysis (computer program) 

Flight simulator for advanced aircraft 

Flight Simulation Laboratory 

Forward 

General purpose computer 

Heading alignment cylinder 

Heading situation indicator 

Head-up display 

Hertz 

Instantaneous center of rotation for elevator inputs 

Inertial measurement unit 

Johnson Space Center 

Knots equivalent airspeed 

Landing and Approach Higher Order System 

Lower order equivalent system 

Langley Research Center 

Modified maximum likelihood estimator 

Master products data base 

Non-intrusive parameter identification program 

National Aerospace Laboratory (the Netherlands) 

Orbiter experiment 

Orbiter Flight Test 

Flying qualities and flight control system design 

criteria experiment 

Operational instrumentation 

xv 
PCM 

PIO 

PTI 

RCA 

RHC 

RSS 

SMS 

STA 

STI 

STS 

TAEM 

TIFS 

TRFN 

USAM 

VMS 

TR-1197-1 

Pulse code modulation 

Pilot induced oscillation 

Programmed test input 

Rate gyro assembly 

Rotational hand controller 

Relaxed static stability 

Shuttle Mission Simulator 

Shuttle training aircraft 

Systems Technology, Inc. 

Space Transporation System 

Terminal area energy management 

Total Infllght Simulator 

STI transfer function computer program 

STI unified servo analysis method computer program 

Vertical Motion Simulator 

xvi 
SECTION I 

INTRODUCTION 

This study is a continuation of work initiated under Contracts 

NAS4-2834 and NAS4-2940, Space Shuttle Flying Qualities and Flight Control System Criteria Assessment, Phases I and II. It provides continuation and refinement of a program for the Orbiter Experimental Program 

(OEX) titled Flying Qualities and Flight Control Systems Design Criteria 

Experiment (OFQ). 

Flying qualities criteria for advanced aircraft have been based on 

many years of experience with civil and military aircraft. For evolutionary designs this experience has provided an orderly and continuous 

base of data that could be applied to each new design with a modest 

extrapolation. However, the Space Shuttle combines the characteristics 

of a spacecraft and aircraft. It is radically different in configuration, operational envelope, and complexity than any vehicle flown 

before. It is a highly augmented, fly-by-wire vehicle whose control 

system design preceded by several years those of current military aircraft. Consequently, large extrapolations had to be made to establish 

handling qualities and flight control system design criteria for the 

atmospheric flight phases of the Shuttle mission. These criteria are 

based primarily on Shuttle-specific simulations and on experience with 

high performance aircraft; however, because the Space Shuttle is a large 

departure from pas t experience, much uncertainty has exis ted as to the 

validity and application of existing criteria. 

The purpose of this continuing effort is to define an effective program of flying qualities and flight control system design criteria 

experiments (OFQ). l'he first phase effort, documented in Ref. 1, was 

devoted to review of existing flying quality and flight control system 

specification and criteria; review of Shuttle experimental and flight 

data; identification of specification shortcomings; and preparation of a 

. preliminary OEX approach to produce the optimum use of flight data to 

develop modified flying qualities criteria for Space Shuttle craft in 

general. 

TR-1l97-1 1 
The Phase I investigation identified several likely problem areas to 

be addressed in the OEX plan. First, mismatch of Shuttle specification 

pitch rate response boundaries (and Shuttle response) with available 

flying qualities data raised the question of whether the specification 

response boundaries are misplaced or whether the available data base is 

inadequate for highly augmented relaxed static stability aircraft. The 

specification boundaries also appeared to allow excessive pitch and roll 

rate response dead time. Second, comparison of Shuttle characteristics 

with other criteria, guides, etc., tended to indicate it exhibited 

excessi ve longitudinal and lateral effective time delays. This would 

lower the effective vehicle bandwidth and then reduce pilot-vehicle and 

autopilot-vehicle attainable closed-loop bandwidth in rolling and path 

control functions. It would also be expected to produce a tendency for 

PIO under high stress, precise control situations. Other likely problem 

areas concerned pilot location effects and hand controller characteristics. While well ahead of the c.g., the pilot is aft of the center of 

instantaneous rotation for longitudinal control inputs. This location 

has consequences on longitudinal path control (possibly quite unfavorable for precise control situation) and lateral acceleration at the 

pilot, station. The rotational hand controller (RHC) displacement/ 

force/electrical command, combined characteristics possibly result in 

larger pilot control latencies (due to near isotonic properties). This 

can also effect the control bandwidth and contribute to control difficulties in urgent tasks. Finally, the Phase I effort indicated possible 

problems concerning off nominal cases of critical aerodynamics variation 

sets, trim extremes, and reduced surface rates. 

The Phase II investigation (Ref. 2) continued the review and analysis of applicable experimental and Shuttle flight data and provided further definition of the Orbiter Flying Qualities Experiment (OFQ) Plan. 

In particular, the influence of "superaugmentation" on vehicle handling 

characteristics was continued. It was found that the Shuttle qualifies 

as a superaugrnented vehicle and the Shuttle 

appropriate for 

specification pitch rate 

response this class of vehicle. 

Further, 

TR-1197-1 

boundaries may 

superaugmented 

be 

aircraft have unconventional attitude/path 

2 
response characteristics and lack speed stability. Review of STS-1-4 

crew qualitative assessments indicated flying qualities to be adequate 

at high altitude and speed but support the unconventional, and possibly 

marginal, flying qualities in terminal control (preflare, shallow glide, 

and final flare). 

The OFQ plan was further refined to address the above superaugmentation considerations and special conditions of Shuttle flights which 

require a somewhat unconventional, indirect, experimentation approach. 

The indirect approach consists of inflight experiments combined with a 

correlated research simulati.on program. The unconventional features of 

the approach is the use of non-intrusive flight measurements for effective vehicle and pilot strategy (model) identification. These flight 

measurements are then used to validate simulations for ground experiment 

programs involving critical flight situations not likely to be permitted 

(or encountered) on Shuttle flights. 

The purposes of this Phase III study encompass continued analysis of 

Shuttle flight data, conduct of supportive analytic and simulation 

efforts, and further refinement of the OFQ plan. A major portion of the 

effort has been devoted to exploring the crucial, but unperfected, 

flight data measurement and reduction techniques required for the nonintrusive experiment approach. This is documented in Section II which 

encompasses an overview of currently available flight data, identification of the effective augmented vehicle from the flight data, pilot 

technique identification and modeling, and, finally, a summary of flight 

data problems and recommendations for the OFQ experiment. 

Section III continues the Ref. 2 summary of crew qualitative assessments and presents comments on the STS-5 and 6 entry flying characteristics gleaned from systems debriefings and associated press releases. A 

second subsection outlines considerations which led to a revised questionnaire for crew in-depth debriefing. 

presented in Appendix B. 

The revised questionnaire is 

A joint STI/DFRF preliminary experiment performed to quantify the 

influence of Shuttle RHC characteristics on pilot latency is described 

TI{-1l97-1 3 
in Section IV. Results of this simple experiment lead to the conclusion 

that the Shuttle RHC configuration and force/displacement characteristics do increase the pilot's neuromuscular time delay. 

Overall conclusions and OFQ recommendations are summarized in 

Section V. 

TR-1197-1 4 
SECTION II 

PILOT/VEHICLE IDENTIFICATION FROM FLIGHT DATA 

A. OVERVIEW OF FLIGHT DATA ANALYSIS 

1. Flight Data Analysis Goals 

Three primary goals guided the analysis of flight data in Phase III: 

• Identification of the Effective Augmented Vehicle 

While there has been extensive effort to identify the 

aerodynamic coefficients of the Shuttle airframe from 

flight data, there appears to have been no effort to 

similarly identify the effective vehicle, as seen by 

the pilot, which is dominated by the flight control 

system. This activity is important to verify analytical models developed in Phase II (e.g., the superaugmented pitch response). Furthermore, a well 

defined controlled element model is necessary for the 

second goal -- pilot technique identification. 

• Identification of Piloting Technique 

Identification of piloting technique is more difficuI t because of technique variations among pilots, 

pilot remnant and uncertainty about cues. 

• Refinement of OFQ Procedures 

In addition to obta:i.ning quantitative definition of 

the vehicle and pilot, refinement and verification of 

the identification procedures proposed in Phase II 

was a primary concern because of the emphasis on nonintrusive techniques. 

The OFQ plan as developed in Phase II placed first priority on the 

approach and landing, from the terminati.on of the HAC turn through 

touchdown (see Fig. 1). The Phase III flight data analysis has involved 

all flight segments except the steep gli.de. 

2. Shuttle Flight Data Sources 

The primary source of data for the Phase III flight data analysis 

were the computer files available at DFRF for use with the Modi.fied 

TR-1197-1 5 
rJ') 

~ 

w 

!iii: 

'0 

., -

., 

Q. 

~ 

.. 

4 

.... .) c: 

Cl 

:l>' :i 

cr 

lAl 

300 

250 

200 

150 

.J 

(!) 

<[ 

b) Equivalent Airspeed Variation 

-3060 It 

I 

At- 2893211 '" ~ ",'" 'I ,/ I 

",'" II 

",/ ......... ,9°-.u 

,/,/ I I 

/' 'I ~\,// I I 

'0'00/ I I ~.;:/ I 

'" I I ° / --J r-1.5 

"" I I 

/"" I 

// I 

- -170011 I I :: I 

I gl 

I ~I 

I "I I 0: I 

I I 

I 

Not to Scale 

Runway 

Threshold 

X=o 

Shallow 

Glideslop 

Aim Point 

1000 It 

Touchdown 

! hI =5011-

I Steep Glide I Equiliburium 1 ..... ---- Prellore Pullup -LJ Sh~IIow Final 

Slope Glide on Glide Flare 

Capture Steep Glideslope Exponential 

TR-1197-1 

Flare to 

Shallow Glideslope 

aj Trajectory 

Figure 1. Shuttle Approach and Landing 

(Nominal Characteristics) 

6 

------ Down Range 

Position 
Maximum Likelihood Es timator (MMLE) program. These files exis t for all 

Shuttle flights to date, and have been used extensively for extraction 

of airframe aerodynamic coefficients. These files may be made available 

as local files on the DFRF Cyber computer and have been transmitted over 

telephone lines to the STI PDP-ll computer. Significant effort was 

required to develop this data transfer procedure; but it is thought to 

be considerably cheaper, faster, and simpler than physically transferring magnetic tapes. While the MMLE files were set up for airframe 

aerodynamic identification, they are usable for preliminary Flight Control System (FCS) and flying qualities studies. However, it will be 

necessary to augment the data available and resolve certain problems 

before further progress can be made. Specific problems and data needs 

will be detailed in Subsection D. 

Two other sources of flight data were also investigated: the Master 

Products Data Base (MPDB) at NASA JSC and the Best Estimated Trajectories (BET) computed at NASA LRC. To date, it has not been possible to 

use these sources. The problems and suggested solutions will also be 

discussed in Subsection D. 

Efforts to obtain the STS-4 data from the DFRF MMLE file were begun 

in Phase II. However, transfer of the data over phone lines to the STI 

computer proved more difficult than expected and was not completed until 

the Phase III effort had begun. It was decided to continue with the 

ST8-4 data since use of data from any later flight would require additional effort and no unusual events occurred in the 8TS-5 flight -- the 

only a.dditional flight at the start of the analysis. Also, the most 

relevant pilot commentary is available for STS-4. 

3. Attitude Control Time Histories 

Five time histories: Rotational Hand Controller (RHC) deflections, 

pitch rate, angle-of-attack, pitch attitude, and normal acceleration are 

shown in Fig. 2 for the preflare through touchdown region of the 8T5-4 

landing. These traces include the variables pertinent to inner loop 

(attitude) control. To consider outer loop (path) control, sinkrate and 

altitude are required; however, because of data complications this is 

TR-1l9l-'1 7 
NOSE 

UP 

8 RHC 

(deg) 

NOSE 

UP 

Q 

5.0 

2.5 

0 

1.0 

...... :----- Pref I are 

Rotational 

Hand 

Controller 

I Pitch Rate 1 

Touchdown 

Shallow Final 

----__..-t-a1E- G I ide -....-.-- Flare-~ 

Termination 

Flare 

Initiation 

(deg/sec) ~ O~=---------------------------H~~~~----~-----~~=-~~~~~ 

a 

(deg) 

e 

(deg) 

nz 

(g'5) 

I Angle of Attack I 

5 

0 

e 

10 

5 

15t 

1.0 

~ I 

0 5 

I 

10 15 20 25 

Time (sec) 

Figure 2. Attitude Control Time Histories 

STS-4 

TR-1197-1 8 

e 

a 

30 35 
deferred unt il Subsection C. The partitioning of the time his tories 

into preflare, shallow glide, and final flare segments is based, in 

part, on direct examination of the time histories shown; but also on 

additional analyses such as the altitude/sinkrate phase plane analysis 

to be discussed in Subsection D. The transitions between preflare and 

shallow glide, and between the shallow glide and final flare, may each 

be associated Hith a discrete RHC pulse (crosshatched in Fig. 2). 

Indications of piloting technique variations among the flight segments may be seen from the traces; in particular the RHC trace. During 

the preflare maneuver, the RHC trace indicates an initial applied acceleration command followed by a higher command level. The transition from 

the steep glide to the preflare resulted in a near-neutral stability 

oscilla.tion in the pilot/vehicle system with a natural frequency of 

approximately 2 rad/sec. This oscillation will be considered further in 

Subsect.ion C. 

RHG input and pitch response in the preflare task indicates fairly 

continuous pilot control action. However, in the shalloH glide and 

final flare, the RHC input has a different character and consists of a 

series of pulses which are either positive or negative and (unlike Q) do 

not pass through zero. This activity ceased for slightly less than a 

second before and after the discrete nose-up pulse which initiated the 

flare, then a pulsive tracking activity continued from preflare to 

touchdown. Pulsive pilot output is often associated with difficult controlled elements requiring low frequency lead (Refs. 3 and 4) and is 

currently being investigated for superaugmented configurations in a 

related study (Ref. 5). 

B. IDENTIFICATION OF THE EFFECTIVE AUGMENTED 

VEIllICLE FROM FLIGHT DATA 

Because of the dominant effect of the FCS on the Shuttle's pitch 

response, identification of airframe characteritics alone is not adequate to characterize the effective vehicle seen by the pilot. The 

superaugmentation model was developed in Phases I and II to deal with 

this situation. In Phase III an effort has been made to verify this 

TR-1l97-1 9 
model from Shut tIe flight data. This section will begin with a revi,ew 

of superaugmentation theory to examine the Low Order Equivalent System 

(LOES) form implied for the pitch rate transfer function q/qc and also 

to consider artifacts of the approximation which may contribute to anomalies in identification. With this theoretical background, identification of the Shuttle q/ORHC transfer function using spectral methods and 

the Nonintrusive Parameter Identification Program (NIPIP), will be discussed. 

1. The Superaugmentation Model 

The simplified block diagram (Fig. 3a) shows a representative pitch 

rate based system appropriate to superaugmented aircraft such as the 

Shuttle. The basic fedback loop consists of pitch rate measured with a 

rate gyro fedback to equalization consisting of a lead with time constant Tq and a lag which is a pure integrator. In the Space Shuttle 

this equalization is created by a oe feedback through a first order lag 

around the elevator servo (Ref. 1). The forward loop contains the actuator, bending mode filters, and other elements which contribute to 

effective high frequency lag in the system. Provision is also made for 

turn compensation so that the pilot is not required to maintain RHC 

pitch deflection in a steady turn. 

Perhaps the best way to view the superaugmentation approximation is 

through the use of the Bode root locus plot. The Bode magnitude plot is 

sketched in Fig. 3b for the open loop q/qe transfer function of a nearneutral airframe. This plot is particularly useful for visualizing the 

important relationship between the dynamics of a relaxed static stability (RSS) airframe (e.g., the Shuttle) and a superaugmented flight control system. The effects of RSS are evidenced by the real short period 

poles I/TsPI (stable) and 1/TsP2 (unstable). The phugoid appears somewhat more conventionally as a low frequency complex pair and the airframe attitude zeroes l/Tel and I/Te2 are relatively uneffected by 

relaxation of static stability. 

The effects of the flight control system design are contained in the 

equalization zero I/T and the asymptotic crossover frequency Wc which 

q a 

TR-1l97-1 10 
Q o = Ro ton po 

:::~~,'.]: R 

Pilching 

Pilot 

comn~~tic S" Filt '-p 

- Velocity Equalization Ac tuator - Error Elevator k 

~ce and Aircraft 

er qe 

Kq(s+ I/Tq ) 

Bending Mode 8 ----- Dynamics 

- s Filters 

Pitch Rate .. 

a) Typical Pitch SAS For Superaugmented Aircraft 

I (Iosed Loop Root 

IGII"")! and IS. GIl"') 

-- CI0scd LOop 

Yawing 

Velocity. R 

Bonk 

Angle, p 

Pitching 

Velocity,q 

IGIJwl1 and <l.G!]w} Open 

- - -- Loop, Asymp10te .73 1.06 

_ !GI-a)! LeI! HOlf Plene 

Istoble) Root Locvs 

IGkr11 Right Half Plane 

---- lunsroLICl ROOI LoCus 

••••••••••• , CO"'pl& ~ Plane Root Locus 

OdB for ~e/ected Kq 

.44 

I 

Tq 

-1290PSfl 

135 psf dB 

= I. 5 rod /sec 

b) Bode Root Locus Sketch, Superaugmented FCS 

Shuttle Pre flare To Touchdown 

Figure 3. Superaugmentation Concept 

TR-1197-1 11 
is set by the selection of the loop gain Kq • It may be seen that the 

crossover region is defined primarily by the FCS parameters. Thus, to a 

first approximation, the dominant closed loop root is determined by 

l/Tq and wc

a

' Asymptotic approximations in the crossover region lead to 

the superaugmentation pitch attitude relation (Ref. 2): 

L K(l/Tq)e -TS 

(la) 

qc [l;, wn ] 

l; I "2 ITqWca 

(l b) 

wn Iwc/Tq (lc) 

K w~ITq (ld) 

The adequacy of these approximations is determined by the behavior 

of the airframe roots as the pitch rate loop is closed which may be seen 

in the root loci (heavy lines) in Fig. 3b. The unstable short period 

root I/TsP2 migrates from the right half plane through the origin and 

into the left half plane to approach l/TSI and form an approximately 

canceling dipole. The stable short period root I/TsPl is driven into 

1 ITs 2; a situation which will occur whenever l/Tq > l/TsPI ' The adequacy of the superaugmentation approximation shown in Eqs. la-d) is 

dependent on the extent to which l/T;PI and I/TS2 cancel. The behavior 

of this dipole, which is fundamental to the superaugmentation concept, 

is determined by basic RSS airframe characteristics. 

The dimensional aerodynamic coefficients which determine the values 

of l/TsPl and 1/TS2 are functions of dynamic pressure q which decreases 

from 290 to 135 psf from preflare through touchdown. However, they are 

essentially independent of Mach number, which varies from 0.48 to 0.31 

in this region of essentially incompressible flow. The high frequency 

attitude numerator I/TS2 may be approximated as 

TR-1197-1 12 
= (4) 

This approximation is generally adequate for conventional aircraft 

and even better for RSS aircraft with low values of Mw. Based on this 

approximation, I/Te2 decreases from approximately 0.64 to 0.44 rad/sec 

at touchdown as indicated in Fig. 3b. 

Estimation of the stable short period root requires knowledge of the 

Shuttle static margin which may be obtained from a 'CL - CM' plot such 

as shown for typical conditions in Fig. 4. Flexibility effects are negligible in this region and the static margin is 

3.7%c unstable 

The maneuver margin is: 

= 0.9%c unstable 

For an airframe with near neutral maneuver margin, the dipole ratio 

(1 ITsp 1) I (lITe 2) may be developed in terms of CM

q 

and CLa as 

I/TsPl 

1 Mq 

1/Te2 = +-

Zw 

1 1 (C )2CMq = -Z Ky CLa 

1.67 (Shuttle) (2) 

Thus the pole and zero maintain a constant and fairly close rela-

~ionship throughout the region of interest. Several conclusions may be 

drawn regarding the dipole's effect. The upper end of the airframe 

TR-1197-1 13 
MACH I\!O. 

0.400 

SYM O<=:UA E 

A -35.00 

8 -30.00 

C -20.00 

o -10.00 1.4 

E -5.00 

F 0.0 

G 5.00 

DEL 8F 

-11.700 

Cj=o 

DEL S8 

25.000 

Q8AR1 

--- q = 300 psf 

XCG (IN) ZCG (IN) 

1098.530 375.000 

= SOLID QBAR2 = DASH 

QBARL-PSF 

0.0 

08.,\R2-PSF 

300.000 

H 10.00 1.2 ~--+---t----+---t-----~7 

17.50 I 

15.00 

I 15.00 

J 20.00 

...J 

U 

·11 

Z 

U 

fZ 

W 

U 

..... 

lJ... 

lJ... 

w 

0 

u 

w 

u 

0:: 

0 

lJ... 

-l 

<: 

L: 

0:: 

0 

Z 

O.S 

0.4 

0.2 

a 

-0.2 

-0.4 

-0.6 

-O.S 

-1.0 I I I I 

.30 

, ! I I I I I I I ; I I I I I 

.25 .20 .15 .10 .05 o -.05 -.10 

PITCHING MOMENT COEFFICIENT ( 0.5570 L8) 

Figure 4. Shuttle Airframe Longitudinal Stability, 

Flexible-Approach and Landing 

(Adapted from Ref. 6) 

TR-1197-1 14 

12.50 

5.00 

I - O. OO=ALf5:c;!A 

I : I j' I I 

-.15 -.20 
dynamics region is set by the high value of 1/TsP1 ' which is about 

1 rad/sec and thus less than l/Tq = 1.5 rad/sec for the Shuttle. The 

pole and zero have a constant ratio of 1.67, thus even without closure 

of the q + 0e loop there is a tendency for dipole cancellation which is 

increased when the loop is closed. It should be noted for later reference that on the Shuttle there is a first order filter (the 'ELERROR' 

filter) in the region of the high frequency dipole. The filter zero is 

near 1/T62 and the filter pole is close to 1/TsP1 ' thus producing a 

lead-Ia.g which increases the phase margin in the crossover region. 

The airframe effects are thus limited to higher order artifacts in 

the superagumented pitch response which is effectively determined by the 

characteristics of the FCS as implied by Eqs. la-d. For the Shuttle, 

l/Tq is set by the scheduling (GTRE_COMP) in the 'ELFBK' filter as 

= 1.5 rad/sec M t\ 3 

The variation of the crossover frequency is more involved and is the 

one area where an airframe characteristic, elevator effectiveness, has a 

direct effect. The asymptotic crossover frequency is: 

(3) 

for the Shuttle situation in which wC

a 

> l/Tq • The scheduling of Kq is 

more complex than for Tq and provides an opportunity to examine the gain 

scheduling problem for one of the few operational superaugmented aircraft. The basic scheduling (GD~COMP) is shown in the block diagram of 

Fig. 5 and indicates that 

f(M) Kq IX ---. 

/q 

(4) 

It should be noted that none of the relevant nondimensional aerodynamic 

coefficients are functions of Mach number in this region, thus the 

TR-1197-1 15 
4 (pst) 

-

q 

( pst) -t ~ 

a = KPIT CSSon/y b 

16.0 I 

I 

Kq I a I a 

12.5 I - -----,----" I I , I b (GDQ) I " I I I 

0 I : ' I , , , 

.3 .45 1.2 

M 

Figure 5. Schedule for Pitch Rate Loop Gain, 

Kq (GDQ_COHP) 

rationale for the l'1ach number schedule is of interest. At these low 

altitudes (below 2,000 ft) where density and acoustic velocity are 

roughly cons tant, the Hach number schedule effectively performs the 

function of an additional dynamic pressure schedule. When approximate 

calculations of this effect are made the resulting crossover frequency 

schedule is 

0.177(/q - 0.026 q) rad/sec (5) 

Figure 6 shows this Wc variation from the start of preflare through a --

touchdown, compared to a fixed crossover schedule (Le., Kq <X q-I) and a 

- crossover frequency schedule proportional to the square root of q. In 

the flights of the Approach and Landing Test (ALT vehicle) Kq was scheduled inversely proportional to dynamic pressure Ref. 7, p. 245 thus giving an effectively constant crossover frequency. This has been modified 

to the present OFT schedule primarily because of problems with elevon 

saturation at low speeds with the original schedule. 

TR-1l97-1 16 
2.0 

;:>0, 

Ut 

c: 

Q;I 

:::1 1.5 

Fixed Crossover, Kq'" I/q c:r 

Q.l 

.... 

ll.. r-

.... (,) (1) Q) 

:::- In () ........ III 

II) -0 

C) 0 ... ... u 

(.) C 

+- U 

() 3 ... -

().. 

E 

>, 

~~ 

« 

1.0 

OL--~ 

o 

Touchdown 

[css Mode Onl~ 

200 

q (psf) 

Dynamic Pressure 

Figure 6. Effective Pitch Gain Scheduling of Shuttle 

2. Possible Methods and Inputs for 

Effective Vehicle Identification 

Start of 

Preflore 

300 

Several identification methods for the effective vehicle are available. First, spectral analysis methods may be used, in particular the 

use of the fast Fourier transform implemented in programs such as the 

STI Frequency Domain Analysis program (FREDA), Ref. 8. Secondly, time 

domain least squares estimation routines are available, such as the STI 

Non-Intrusive Parameter Identification Procedure (NIPIP), Ref. 9. There 

are, :Ln addition, the airframe parameter identification programs (e.g., 

~~LE) extended to augmented vehicles by various means. One approach has 

TR-1l97-1 17 
been to assume a conventional linear aircraft model and identify effective stability derivatives; however, this approach has the disadvantage 

of not identifying unusual effective derivatives (e.g., Me) unless these 

effective derivatives are accounted for in the original model form. 

This problem may be avoided by structuring the model in terms of the 

effective poles and zeros of the vehicle response. Some rational basis 

(e.g., superaugmentation theory) for selecting the number of and constraints on poles and zeros would still be required. 

All of the above identification methods require some form of input 

to the system and there are a number of possibilities. Manual pilot 

activity actompanying normal flight operations is the simplest, and in 

the case of the Shuttle OFQ may be the only pilot input available. A 

sum of sine waves, on the other hand, has theoretical properties which 

are particularly advantageous for obtaining good signal-to-noise ratios. 

The frequency sweep, which begins with a low frequency quasi-sinusoidal 

input and increases to higher and higher frequencies, is particularly 

useful because it is easily generated by pilots and has a reasonably 

flat power spectrum. Finally, pulse train inputs can be tailored to 

have desirable power spectral characteristics. For the OFQ it may ultimately be possible to tailor existing Program Test Inputs (PTI), 

presently used for airframe identification, to form an approximation of 

a sum of sine waves disturbance for pilot tracking experiments. 

For the Phase III work, both the FREDA and NIPIP procedures have 

been used on data with normal pilot inputs and will be discussed in the 

following sections. 

3. Identification of Superaugmented 

q/6RHC Using Spectral Methods 

The spectral identification of the q/oRHC describing function 

requires the 0RHC and q time series (Fig. 2). The FREDA program obtains 

the spectral density distributions ~oo and ~qq and cross spectral ~qo by 

direct Fourier transform of the time series using the Wiener-Khinchin 

relationship, Ref. 10. The q/oRHC describing function is then given by 

O~c(jW) ( 6) 

TR-1l97-1 18 
The FREDA program produces discrete magnitude and phase angle pairs 

for q/oRHC, ~oo and ~qq' A coherence function 

2 

w) (7) 

is also computed and gives a measure of the degree to which the output 

is linearly correlated with the input. A p2 of zero implies no correlation and a p2 of 1 indicates perfect correlation between output and 

input u For vehicle dynamic identification, p2 values between 0.8 and 

0.9 are generally indicative of meaningful identification. 

Implicit in the use of the spectral procedure is the assumption that 

the describing function will be time invariant. Because of the properties of superaugmentation this will be true to a first approximation for 

q/oRHC in the region of interest given the fixed value of Tq and the 

small variation of Wc with q (Fig. 6). This situation would not occur 

a 

for an airfame transfer function such as q/oe and special accommodation 

would be required. There are in addition system nonlinearities which 

could, in principal, compromise the use of FREDA; in particular the PIOS 

filter and the stick shaping (ESHAPE). During the STS-4 landing, the 

PIOS filter was active only when the commander executed the preflare 

termination. To examine possible PIOS effects, FREDA runs were made 

over the entire period from the start of preflare to touchdown and for 

subsegments excluding the PIOS activity. The above .considerations of 

time varying system parameters and occasional nonlinear events (e.g., 

PIOS activity) imply a desire for a short identification run length 

TRUN' On the other hand, Ilk1.ximizing. run length is desirable to obtain 

good low frequency data since the theoretical lower bound on the frequency response wmin is 

Wmin (8) 

The FREDA output plots for maximum feasible run length (start of 

preflare to touchdown, TRUN = 30 sec, Wmin = 0.2 rad/sec) are shown in 

Fig. 7. The coherence p2 values are above 0.8 out to approximately 

TR-1197-1 19 
o.~o ~.oo w (rod/sec) 

:20 -.J r 

217' I TRUN = 30 sec I I-q (jwll 0 TRUN 

ORHC dB 

X X X X 

X -:20 _ X 

0 - x 

~(jwl x 

4 X RHC X -~oo 

x 

(degl x 

-:200 - Transfer Function 

o.~o ~.oo w (rad/secl 

:20 -.J 

I ct>881 dB 

D 

I <Pqq IdB 

-:20 _ 

1 -

p2 

D -

TR-1197 -1 

I 

:r: 

0 :r: <PSS 

I I I 

0 :x: 

0 

D 

0 

Power Spectral Density <Pqq 

D 

-

Coherence 

Figure 7. FREDA Output for q/oRHC 

STS-4 Preflare Through Touchdown 

20 

~o.oo 

I 

p limit 

+ 

+ + 

x ++ + + + J 

J j J 

J J+ 

J 

J 

J 

x J 

~O.OD 

I 

I 

1: 

I 

I 

I 

0 I 

J 

II 

Il::o: D 0 

o 0 0 

o 00 9J 

- ~- - --
10 rad/sec. Above this frequency the coherence decreases and thus 

10 rad/sec is taken as limit of validity for the frequency response (the 

describing function plot symbols are changes when p2 drops below 0.8). 

This is a very desirable situation which has important implications for 

the use of non-intrusive procedures in the OFQ, since it indicates that 

normal pilot RHC activity may be adequate for identification. An indicatioa of the identification potential of the RHC input may be seen in 

the relatively flat input power spectra ~oo for 1. < w < 10. rad/sec. 

Some feeling for this spectra may be obtained by examining the 0RHC 

trace in Fig. 2 and noting the low frequency trends, especially in the 

preflare region, combined pulse-like inputs with high frequency content 

in the shallow glide and final flare which effectively provide a rough 

approximation to a frequency sweep. 

Figure 8 presents the corresponding FREDA output for the shallow 

glide and final flare starting just after the PIOS activity and terminating just before touchdown. The results are quite consistent with 

the Fig. 7 data except for the reduction in low frequency data due to 

the shorter run length (TRUN = 12.0 sec). These results are valuable 

because they imply usable results may be obtained for TRUN on the order 

of a flight segment length and they confirm the approximate time invariance of q/ 0RlfC. 

The FREDA program defines the q/oRHC describing function as a set of 

discrete magnitude and phase angle points. Definition of the specific 

parameter values in the superaugmented response form (i.e., the poles 

and zeroes) requires 'fitting' this form to the FREDA output. This has 

been done with the Multi-Frequency Parameter identification program 

(MFP) which provides a weighted least squares fit to the specified 

response form. For fitting a superaugmented q/ORHC form, the MFP program was set up to minimize C (see Fig. 9) where 

C (9) 

TR-1l97-1 21 
l..00 w (rad/sec) l.0.0[J 

I I 

2.". 

I TRUN = 12.0 sec I 

TRuN p limit 

x x 

~ x 

x 

-q-(jw) [J 

8RHC dB 

-2[J 

j + 

J 

[J -

X 

4 rOw) -l.[J[J - X 

RHC x 

(deg) x 

-2[J[J Transfer Function -

[J.l.[J l..[JC w(rad/sec) l.[J.[J[J 

2[J -.J I I 

1<P881dB 

[] 

:I :I c;P88 

lcI>qq/dB 

:I :I 

[J [J J: -2[J _ J: 

[J 

[J 

Power Spectral Density c;Pqq IJ 

[J 

1..[J _ 

~ p2 

Coherence 

[J _ -

Figure 8. FREDA Output for q/CRHC, STS-4 

Shallow Glide and Final Flare 

TR-1l97-1 22 

+ 

+ 

+ + 

J t 

+ J J 

J 

J 

:I 

:I 

J: 

J: J: 

[J J: 

[J [J [J [] IJ 

- --
GFR{ j W k) - GSA ( j W k ) 

,-----lI-.-·Re 

Figure 9. Definition. of Cost Function. in MFP Program 

where 

Wk is the kth matching frequency 

wk is the kth weighting factor 

K(1/T )e-TS 

---q--- [ 1:, wnl (superaugmented form) 

(from FREDA) 

The weighting factors were set to 

(10) 

which essentially imposes the same penalty in dB on magnitude misma.tches 

at all frequencies. The fitting operation has been perfonned both with 

tile lead l/Tq fixed at the Shuttle FCS value Cl/Tq = 1.5 rad/sec) and 

also with the lITq free. The two results are shown in Fig. lOa. It may 

be seen that the two cases are very similar and that both provide a very 

satisfactory fit. A time domain comparison (Fig. lOb) leads to a similar conclusion. 

Table 1 shows a comparison between superaugmentation theory and the 

values extracted from the STS-4 flight data (both for fixed and free 

TR-1l97-1 23 
0.10 

0--1 

jr(jW)1 

RHC dB 

-20 -

l} ~(jW) ORHC 

-100 -'-

(deg) 

-200 -

/ 

1.00 w (rad/sec) 

I 

x X X X X FREDA Output 

F· d I 5 -I Ixe 1" = I., sec q 

q .592(1.5) -.156. -= e ORHC [.74,1.68] 

---- All parameters free 

q .6060.03) -.159. 

ORHC = [.77,1.44] e 

a) Frequency Response Of Filled Superaugmentation 

Forms Compared To FREDA Output 

VI I VI 1. Free 0" 

"- q 0" 1.5 . 

Q) 

(; 

a: 

..c: 

!::! 1.0 u: 

"0 

Q) 

N 

-0 

E .5 IF NLR Boundaries (Ref. II) 

(; 77T" 1977 STS Boundaries (Ref.12) z 

00 2 3 4 5 6 7 

Time (sec) 

b) Normalized P/tch Rate Response To A 8RHC Step 

Figure 10. Comparison of Flight-Derived Effective q/oRHC 

with l/Tq Fixed and Free 

TR-1197-1 24 

10.00 

I 
TABLE 1. COMPARISON OF PARAMETER EXTRACTED FROM ST8-4 

FLIGHT DATA WITH THE 8UPERAUGMENTATION MODEL 

SUPERAUGMENTED EXTRACTED FROM STS-4 FLIGHT 

PARA11ETER _._-_. __ ._ .... _-_ .. _--_. __ ._-- -- MODEL 

l/Tq 1.5 r/s t/Tq = FREE 

-

q/<SRHC(O) 0.17 0.31 0.30 (rad/rad/sec) 

--- .. -.--.---------.----

1/Tq (rad/sec) 1.5 1.5 1.03 

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

r . 0.5 0.74 0.77 

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

wn 1.5 1.68 1.44 (rad/sec) 

- 1" 0.174 0.156 0.159 (sec) 

numerator time constant) for five q/oRHC parameters. The first line 

comparE~s steady state gain values and shows the largest difference of 

any of the comparisons. The extracted values are approximately 80 percent h:i.gher than the estimated values. While the differences have not 

been resolved, it is presently thought that they may be due to calibration uncertainties in the 0RHC signal. This signal was obtained for the 

DFRF MMLE file from the Backup Flight Control System (BFC8), which does 

not contain all of the forward loop elements of the primary system actually used in the 8TS-4 entry. 

The fixed value of flight derived numerator inverse time constant 

exactly matches the superaugmented model by definition. The extracted 

value with l/Tqf is somewhat lower at 1.03 • However, it is still 

ree 

well above the range of values for 1/T82 in the identification region 

(0.44 to 0.64 rad/sec). The values of the damping ratio obtained from 

flight are approximately 50 percent higher than the superaugmentation 

theoretical prediction. It is presently felt that the primary cause of 

the higher in-flight damping ratio, and perhaps also the lower in-flight 

TR-1197-1 25 
value of l/Tq' is the 'ELERROR' filter. To a first approximation, the 

ELERROR filter increases the open loop phase margin in the crossover 

region which would correspond to the effect of a low value of l/Tq. 

Furthermore, the increased phase margin corresponds to a higher closed 

loop damping ratio as is seen in the flight extracted results. 

The flight extracted values of natural frequency are somewhat higher 

for the fixed Tq case and somewhat lower for the free Tq case with 

respect to the theoretical value of 1.5 rad/sec; however, these should 

be considered in the light of the uncertainty implied by the variation 

of crossover frequency as shown in Fig. 6. Finally, the flight 

extracted values of time delay are actually somewhat lower than the 

value obtained by adding the low frequency phase lag approximates of the 

forward loop elements (Table 2). 

From the standpoint of flight validation of the superaugmentation 

model, probably the most important issue is whether the attitude zero is 

really closer to l/Tq than to l/TS2 • This issue may be addressed by 

considering an alternative fit with the attitude zero constrained to 

1/TS 2 : 0.50 rad/sec. Figure 11 shows such a comparison, in terms of 

asymtotes, between the fits of Fig. 10 and a 'l/TS2 fit' with Wu 

1.5 rad/sec. For this alternative fit there is a significant region of 

TABLE 2. CONTRIBUTION OF FCS ELEMENTS 

TO EFFECTIVE TIME DELAY 

FCS ELEMENT EFFECTIVE TIME DELAY 

-- -

Bending Filter 0.039 

Smoothing Filter 0.039 

Computational Delay 0.046 

Actuator 0.050 

-'- 1----.. -----.. -----

Total 0.174 

TR-1197-1 26 

(SEC) 

.. 

, 
0.1 1.0 w(rad/sec) 10.0 

o 

q -20 ~ (j,w) 

°RHC ASYMP 

dB 

-40 

I 

Attitude Zero (rod/sec) 

FREDA Output 

Fixed J_ = 1.5 

Tq 

I 

xxx 

1 Free -- = 

Tq 

1.03 ------

1 

T~ - 0.50 -.----

Figure 11. Asymptotic Compar.i.son of Iq/OllliCI 

w.i.th Three Oifferent Numerator Constra.i.nts 

I 

+2() dH/dec'lde slope wh.i.ch is inconsistent wi.th the low frequency FREDA 

points. In this connection it should be noted that these FREDA points 

are averages of j 'raw' FFT points and are reliable within the context 

of this comparison. No adjustments of wn would provLde a satisfactory 

, 1/T82' fit .• 

4. Identification of Superaugmented 

q/6 RHC Using the NIPIP Plrogram 

Shuttle vehicle identification has also been performed with the 

NIPIP program (Ref. 9). This program is a running least squares time 

domain estimator which was proposed in Phase II as a primary tool for 

pilot model parameter extraction from Shuttle flight data. However, it 

may also be used for augmented vehicle identification if a proper 

TR-1197-1 27 
vehicle model structure is employed. Use of NIPIP for this purpose had 

two primary purposes. First, previous applications of NIPIP to vehicle 

identification from flight data had been limited and not too successful 

(Ref. 13). This lack of success is thought to have been due to lack of 

adequate resolution in the instrumentation system. A second reason for 

attempting vehicle identification with NIPIP was to test the application 

of this program to actual Shuttle flight data to obtain confidence in 

its use for the more difficult pilot identification process. 

Figure 12 shows the FREDA data (X' s) and the fixed l/Tq fit from 

Fig. 10 compared to results from NIPIP (0' s) • The FREDA run was based 

on the entire region from preflare to touchdown (about 30 sec) and thus 

the low frequency limit (21T/TRUN) is lower than for the 7.3 sec NIPIP 

run. However, in the valid frequency range, the NIPIP result compares 

quite favorably to the FREDA result. 

C. PILOT TECHNIQUE IDENTIFICATION 

A major portion of the Phase III work has been devoted to pilot 

technique identification efforts, both to gain a further understanding 

of pilot technique and to refine identification procedures for use in 

the actual OFQ. From this effort a number of data problems and needs 

have been identified of which altitude and sinkrate problems are the 

most critical (these will be discussed in detail in Subsection D). The 

emphasis in the Phase III activity, consistent with the Phase II OEX 

plan, has been on non-intrusive identification procedures. This is 

based on the belief that severe constraints on off-nominal Shuttle 

flight maneuvers must be expected thus placing high priority on methods 

which can produce useful results from whatever flight data becomes 

available. Three specific analytical procedures will be discussed in 

the following section. The first is a simple analysis of the nearneutral closed loop pilot vehicle oscillation which occurred at the initiation of the STS-4 preflare pullup. A value of effective pilot time 

delay has been extracted from this analysis. The second procedure, the 

altitude/sinkrate phase plane analysis, produced interesting results and 

has promise for the OFQ. This procedure has been applied to the shallow 

TR-l197-1 28 
X FREDA IT = 30 seconds I pref lare through touchdown 

o N I PIP I T = 7.3 seconds I shallow glide 

0.10 

20 --I 

~:!..(jW)1 

Cle dB 

0-

-

-20-

T = 30sec 

~" ~ 

o 

O----_________ .~~ 

o 

4. q~ (jw) 

-100-

(deg) 

-200 -

1.00 w ( rad /sec) 

I 

7.3 sec 

Figure 12. Comparison of NIPIP and FREDA Results 

. for Shuttle Effective Vehicle Identification 

TR-11 97-1 29 

10.00 

I 
glide and final flare maneuver and used to extract parameters of the 

pilot technique model hypothesized in Phase II. This effort contributed 

directly to the final effort, the application of the NIPIP program to 

the shallow glide region. Use of the NIPIP program for pilot model 

definition was not as successful as for vehicle identification. However, some procedural difficulties have been uncovered, some new procedures have been developed, and some data inadequacies have been found. 

1. Effective Pilot Time Delay 

in Preflare Oscillation 

The apparent near-neutral closed loop pilot/vehicle oscillation 

observed at the initiation of the preflare maneuver is shown in Fig. 13. 

As indicated, a period of approximately 3.3 sec is observed corresponding to a natural frequency of 1.9 rad/sec with near zero damping. Under 

the assumption of a closed loop pilot/vehicle system, the product of the 

open loop pilot and vehicle describing functions Yp Yc is theoretically 

-1. Since the Shuttle characteristic Yc has previously been identified, 

the value of Yp (as a complex number) may be computed as 

-1 Yc (1.9j) 

If the pilot is assumed to operate on attitude information 

(0) [0.77, 1.44 J I 

0.606(1.03)e-0.159s s 1 9' • J 

6.50e-0.436j 

If the pilot Yp is assumed to be a pure gain with a time delay Lp 

-0.436 

-1.9 0.23 sec 

TR-1l97-1 30 
Rotatior~al Hand 

. Controller 

NOSE 5.0 

UP 

2.5 

8 RHC 

...... i------ Preflare 

Wo = 1.9 rlsec 

p = 3.3 sec 

(d eg ) 0 '--'---+---t---+--

I Pitch Rote I 115° 

NOSE 

UP 

Q 1.0 

(deg/secl 0 

0 2 4 6 

Time (sec) 

0) Pilot/Vehicle System Response 

Pi lot Shuttle 

Be + r>.. B€ 8RHC 

"QSI" .... Yp Ye 

- ~ 

b) Assumed Pilot/ Vehicle Model 

1m 

Q 

deg/sec 

1.9 

deg 

B 

Figure 13. Elements of Preflare Oscillation Analysis 

TR-1197-1 31 

Re 

6.5~ 8RHC 

deg/deg 
However, we may expect some (first order) pilot lead in the region of 

1-1.5 rad/sec based on the Y

c 

characteristics (Fig. 10). 

accounted for 

Tp 0.7 to 0.8 sec 

If this is 

This is relatively high compared to values usually observed in compensatory tracking, Ref. 3, and may reflect pilot uncertainty in this transition region. 

2. Shallow Glide and Final Flare Pilot Model 

It was concluded in Phase II that landing, and in particular the 

shallow glide and final flare, was the most critical flying qualities 

concern and thus the initial pilot model identification efforts have 

been concentrated on this task. The Phase II activity also produced a 

proposed pilot model for this flight segment which will be briefly 

reviewed and extended here for perspective in the discussion of flight 

data analysis. 

The shallow glide and flare model assumes that the pilot attempts to 

fly a trajectory of the form shown in Fig. 14. The initial conditions 

are set by the altitude and speed (ho and Vo ) at the termination of the 

preflare. The pilot selects a constant value of flight path angle Yo 

for the shallow glide which is maintained to some preselected flare 

altitude hf at which point a flare is initiated in which the pilot 

schedules sinkrate proportional to altitude with time constant Tf • This 

flare law produces an exponential trajectory asymptotically approaching 

a level hB below and parallel to the runway. 

Analytical treatment is complicated by the fact that the Shuttle 

decelerates in this region, but it was shown in Ref. 2 that the deceleration is roughly constant (V ~ 1/4 to 1/3 g) which allowed reasonable 

approximate analysis. The decreasing dynamic pressure does affect the 

vehicle dynamic characteristics (except for q/oRHC), Qut the primary 

effect for the pilot is the creation of a relatively short 'touchdown 

TR-1l97-1 32 
ho 

.c. ,-

Q) 

"tJ 

::J 

-+-

+-

<! 

hf 

Constant Y Exponential 

Shallow Glide Flare 

---- 6xG -- 6Xf 

Ii = VT Yo :: Constant • I 

h=--- (h+hs) Tf 

Touchdown 

Point 

--- V 

Horizontal Distance, x 

figurt~ I,L Idealized Trajectory for the 

Shallow Glide and Plare Trajectory Hodel 

time \vindOlv' to l:leet constraints on touchdown energy, speed, and sinkrate. The model aSSUllles that: the pilot selects h f a priori; thus, as Yo 

is steepened, the glide distance find time are reduced and the speed at 

flar(~ initiation is increased. The hasic equations derived in Ref. 2 

are SUI:lmari2l~d below: 

TR-1197-1 

Speed at flare i.nitiation 

Depth of flare asymptote 

hg 

Touchdown time measured from flare initiation 

t' 

TD 

33 

(lla) 

(11 b) 

(llc) 
Touchdown speed 

Touchdown sinkrate 

Distance traveled in flare 

_ K~ t' 2 

2 TD 

(lId) 

( l1e) 

(11£) 

The effect of the flare time constant is somewhat more complex, but 

it may be examined in the curves of Fig. 15 computed for the nominal 

conditions of Fig. 1 (which are somewhat different than the STS-4 conditions). If the flare is very slow (very large Tf ) there is essentially 

no flare and the trajectory is an extension of the glide. Thus the minimum flare distance is 

(12) 

as Tf + 00 

As Tf is reduced for a faster flare, the trajectory approches a 

level parallel to and hB feet below the runway. When Tf is reduced to 

(13) 

hB goes to zero and the runway is approached asymptotically (AXf + (0). 

For still lower Tf values the runway is never reached and a 'ballooning' 

situation results. Touchdown speed remains fairly constant until 

. * Tf = Tf, at which point the increasing flare time causes considerable 

speed bleedoff. Touchdown sinkrate is strongly affected by the 

trajectory slope, and thus decreases steadily as Tf + Tf. * The Shuttle 

TR-1197-1 34 
10 

8 

u 

<lJ 

Vl 

"-

::: 

a 

.% 6 

2:' 

::: 

.:>:: 

co 

W (f) 

V1 co 

~ 4 0 

-0 

.c 

u 

0 '" f2 

o 

500 10 

400 8-;:: '>-

U X - <lJ 

Vl <J ~ 

~ a 0 

~ 300 6L;: 

-0 

<lJ co 

<lJ 

Cl. -0 

(f) ~ 

<lJ co > ~ 0 

0 

-0 200 4f- L 

u <lJ 

'" u 0 co 

f- 0 

~ 

Ci 

100 2 

o 0 

'>-

-

In 

.c 

<lJ 

0 

0.. 

E 

>-

Vl 

<l: 

~ 

0 

lL 

-0 

.c 

0.. 

<lJ 

0 

250 10 

200 8 

150 6 

100 4 

50 2 

u 

<lJ 

(f) 

Ballooning 

Region 

0 

'..t-

<lJ 

0 

lL 

co 

<lJ 

E 

f-

-0 

<lJ 

(f) 

a. 

0 

W 

'!? 

:> 

"-

.c -

" .-f5 

Vf = 406.1 ft/sec 

K. = 708 ft/sec V' sec 

hrD / 

/ 

/ hTDDesign 

~ Range 

SHUTTLE 

AU TOLAND 

SYSTEM 

10 

Yo" -1.5 deg 

hf = 50 ft 

VrD Main 

Gear Limit ((cce 

eft / / ( 

hTD Main 

Gear Limit 

hrD __ 

..;..;;-- ------=:: / ( if ( ie ( 

lim 

Tf~oo VTD 

);) ))) 

lim ' Tf-oo tTD"-hf/VfYo 

~ r 6Xf (V = constant) 

__ -.i _. 6Xf(K.j =708 fps/sec) 

--------- )/) )/7 ))7 ;; ) 

lim AX' h Tf -00 w. f = - f Iyo 

15 20 

Flare Time Constant, Tf (sec) 

Figure 15. Effect of Pilot's Flare Time Constant on Final Flare 
Autoland flare time const8.nt (Ref. i4) is 5-6 seconds. This places . 

hTu i.n the design range, but unhrtunately also puts the flare irlto the 

sensitive region near the ballooning boundary (T f ,:, TV. 

To this point only path and speed control have been considered but 

these objectives require pitch attitude control. Complete pilot models 

for the shallow glide and final flare were developed in Phase II 

(Figs. 27 and 28, l{ef. 2). These have been combined and sirnplified in 

Phase [[I to the form shown in Fig. i6. The speed control loop in the 

Phase II model was not included, in part, because there were no cle8.r 

indLcations of p<lth moduli:l.tion for VTD control in 8TS-4. The Phase II 

,nodel also provid",d for a feed forward loop to the RHC to accommodate 

precognitive inputs. The STS-4 [{He trace, Fig. 2, i.ndicates that precogni.tive illputs are probably Hmi.ted to pulses at the ElLght segGlerlt 

transitions separated by regLons of closed loop tracking. The feed forward was thus replaced by a discrete Lnput 5 RHC in Fig. 16. The tri:l.nsition from glide to flare is <lccornmodated in the present l:lodel by a 

switch on the sinkrate cOI:lmand. 

Thus tile present model has <l series structure as ex:pected for a CTOL 

technique, especially when pi. lot lead is anticipated in the i.nner loop 

(Ref. i5, pp. 125-142). Some relevant comments are avai.lable from the 

STS-4 commander cr. K. Hattingly) regardin.g the manner in which the . landing is accomplished: "you cannot fly h or h in the Shuttle; you . have to fly pitch attitude. [I] guess h from Hartsfield's altitude 

calls, then move the nose." The intet"pretation of this statement, in 

VTYo 

I/Tf 

hc=-h s 1 

YPh 

flare 

h h 8 

Figure 16. Shallow Glide and Final Flare Pilot Model 

TR-1197-1 36 
light of the Phase II model, is that it is impossible to control altitude or sinkrate directly w:lth elevator, and in fact, an inner pitch 

attitude loop must be maintalned. This is entirely consistent with the 

series loop structure shown in Fig. 16. 

3. Altitude/8inkrate Phase Plane Analysis 

The shallow glide and final flare pilot model of Fig. 16 may be usefully viewed as a trajectory in the altitude/sinkrate phase plane shown 

ideally in Fig 17. If the shallow glide region has constant flight path 

angle Yo as the model implies, the phase plane trajectory will be a 

straight, sloping line. If the sinkrate were more nearly constant, the 

glide trajectory would be horizonta,l. In the final flare region, where 

sinkrate is assumed to be scheduled proportional to altitude, the phase 

plane will be a straight Hne with slope -1/Tf • Ideally, touchdown 

• should occur with the sinkrate in the nominal -hTD range (1.5 to 

2.5 ft/sec). 

Application of phase plane analysis to the 8TS-4 flight was complicated by a number of problems with the available altitude and sinkrate 

data. These problems and their interim (Phase III) solutions will be 

discussed in detail in Subsection D. Briefly, the primary requirement 

to generate a usable phase plane trajectory was generation of a sinkrate 

signal by complementary filte.ring of hand a Z signals. Figure 18 shows . the final form of the STS-4 h-h trajectory. 

The trajectory has been partitioned into final flare, shallow gLide, 

and pre flare region based on its shape and the time histories, shown in 

Fig. 2. The preflare region may be seen to be a region of rap:ldly 

decreasing sinkrate consistent with the pull-up maneuver. This is followed by capture and tracking of the shallow glide slope where the trajectory indicates a lightly damped pilot/vehicle system with a settling 

time comparable to the glide period. Sinkrate in this region varies 

from 5.5 to 4.3 ft/sec (based on the fitted line shown dashed), however, 

when the speed variation is accounted for, the variation in flight path 

angle is smaller, (-0.66 to -0.6 deg). 

TR-1197-1 37 
Touchdown 

.~ 

I 

Q) 

..... 

o 

.... 

..:.c:: 

c 

(f) 

JJ 

o Altitude, h 

Nominal hTD Range 

hoc 

h = constant 

---------~ 

"'-

Yo :: Constant '" 

'\ 

\ 

\ 

Final Shallow Preflare 

Flare--.,-..... ---- Glide -----~~ Pullup ~,-----+--' 

Figure 17. Idealized Altitude/Sinkrate Phase Plane Trajectory 

for the Shallow Glide and Flare Pilot Hodel 

TR-1197-1 38 
Tf (sec) 

2.0 

4.6 

u 

<I> 

VI 

"-

.... 

:::. 

• .s::; 

~ 0 

.... 

.:.: 

c: 

(/) 

5 

10 

15 

20 

2'" ~) 

30 

Final 

Flare --J~I-------- Shallow Glide -----__ t-.--- Pref/are 

Apparent 

Ballooning 

I 

210 

Altitude, h (ft) 

40 50 

~~--~~---~~ . --,...,....,I----,"'>"'7'"-~ Design h at Touchdown 

VT = 410 ft /sec 

Ii = -4.3 ft /sec 

Yo = - 0.60 deg 

STS-4 

70 

t~ ~ ~T = 480ft/sec @ 

h = 5.5 ft /sec 

Yo = -0.66 deg 

Figure 18. Altitude·-Sinkrate Phase Plane Analysis of STS-4 

Shallow Gl ide and Final Flare Complementary Filtered 11 

TR.-1l97-1 39 

90 
The transition from shallow glide to final flare shown in Fig. 18 

corresponds to the RHC flare initiation pulse shown in Fig. 2. The 

local peak of sinkrate seen immediately after flare initiation may be 

traced yo the negative download on the elevons associated with the noseup RHC command. In the flare, sinkrate decreases roughly proportional 

to altitude with a superimposed oscillation which leads to a region of 

apparent ballooning about 5 ft above the runway. It should be noted 

that the altitude scale has been adjusted to give zero altitude at the 

(known) touchdown time to remove bias present in the altitude signal. A 

value for the flare time constant Tf can be extracted by fitting a 

straight line to the final flare region. Because of the oscillatory 

nature of the actual trajectory, various straight lines could be fitted 

as shown in Fig. 18 giving 2 , Tf ' 4.6 sec. If the design ~TD region 

is interpreted as an indication of the desired accuracy of control for 

sinkrate, the implication is that, while the shallow glide and final 

flare performance is acceptable, the manual control precision is less 

than might be desired. Beyond this interesting result for 8T8-4, is the 

important indication that the Phase II pilot model is reasonable, and 

that the parameters Yo and Tf may be simply extracted from the phase 

plane plot, thus confirming the value of the method for the OFQ 

experiment. 

To analyze the pilot's terminal control problem, based on the sha1- . low glide and flare model, contours of XTD , VTD , and hTD were plotted in 

the pilot's Tf - Yo control plane in the Phase II work. These plots 

(Fig. 2, Ref. 2) were constructed based on the simplified flight mechanics resulting of Eq. 11 approximation and the nominal Vo , ho, hf, 

and K~ values shown in Fig. 1 and used in Fig. 15. However, these 

parameter values were considerably different for 8T8-4 as shown in 

Table 3. 

. Revised VTU and hTD plots are shown in Fig. 19 based on the observed 

ST8-4 Vo , ho, hf, and KV values with the 8TS-4 (Fig. 18) values of Tf 

and Yo shown for comparison, are also indicated in Fig. 19. From the 

phase plane trajectory, Fig. 18, the ST8-4 hTD is seen to be near zero 

. with some ballooning indicated. This is consistent with the hTD 

TR-1197-1 40 
TABLE 3. SHALLOw GLIDE A11D FLARE PARAMETERS 

PARAMETER REFERENCE 2 VALUE STS-4 OBSERVED VALUE 

Vo (fps) 468 480 

ho (ft) 150 68 

hf (ft) 50 18 

K· v (fps/sec) 7.08 8.0 

contours in Fig. 19a. The observed VTD was 213 kts which, from Fig. 19b 

would imply somewhat higher (more positive) values of Yo and/or Tf than 

extracted from the phase plane. However, the sensitivity of VTD is very 

high in this region and Yo uncertainties of 0.1 deg or less could 

explain the error. These comparisons are not considered as 'proof' of 

the Fig. 16 pilot model, but simply as indications of consistencies 

between the model and observations. Resolution of basic data problems 

and a.nalysis of data from many more flights- are needed; however, these 

comparisons provide a check on the flight data analysis which will be of 

value for the OFQ. In addition the Fig. 19 plots allow 'what if' questions to be considered about why the STS-4 crew may have chosen their 

strategy and the consequences of alternative choices. There are no 

pilot comments or other explicit indications as to why the STS-4 crew 

flew a low trajectory (Le., ho and hf lower than the Fig. 1 nominal 

values); however, given this situation, the Fig. 19 plots provide some 

poss:Lble explanation of the Yo' Tf choice. Figure 19b indicates that 

acceptable VTD control would be virtually impossible for Yo > -0.8 

deg. This is in marked contrast to the 'nominal' situation as shown in 

Fig. 21b, Ref. 2 and helps explain the apparently low value of Yo for 

STS-li. The STS-4 situation occurs because the low ho limits the time 

. available for speed bleedoff unless a shallow glide is used. The hTD 

situation is not quite so critical, but the ST8-4 Tf margins are reduced 

relative to the Ref. 2 nominal case at a given Yo. The STS-4 Yo ,: 

-0.6 deg produces a relatively large Tf 'sweet spot' but at the expense 

of ballooning potential. 

TR-lJ.97-1 41 
20 

Tf 

(sec) 

10 

y (deg) '0 

b) Touchdown Speed 

(ftlsee) 

hro = 0 -1.5 -2.5 -6 

Tf 

(sec) 

30 hro 

20 

10 

4.6 

2.0 

-9 

-I 

STS 4 

Vo = 480 ftlsee 

ho = 68 ft 

hf =18ft 

K. = 8.00 (ftlsee) 

v (sec) 

Main Gear Limit 

Main Gear Limit, 

Crosswind 

'-----+---'----'------->-,----------------'---

00 -.6 -.66 -I 

STS 4 Observed 

(Tf ,Yo) 

Yo (deg) 

a} Touchdown S/nkrate 

-2 

Figure 19. Variations of Touchdown Variables in the 

Pilot's Tf-Yo Control Plane. 8T8-4 

TR-1l97-1 42 

-3 
4. Pilot Strategy Identification with 

NIPIP in the Shallow Glide 

Efforts to identify pilot model parameters using the NIPIP program 

have been made for the shallow glide region. The pilot model, Fig. 15, . reduces to a h + 8c , e + 0RHC system in the flare. For application of 

the NIPIP program it is necessary to represent the rotational hand con- . troller deflection in terms of the response variables hand e and the 

• external (precognitive) rate command hc ' 

(14a) 

Based on the development in Phase II the expected form of the pilot elements are 

= 

Kh(l/TLi) • 

0!Tli) = 

( l.4b) 

Kh 

s 

where the approximation of Yph as an integrator is made under the 

assumption that the outer loop crossover frequency is well below 1/T 82. 

The result is a five parameter model with unknown parameters: 

One of the important lessons learned from the use of NIPIP in 

Phase III was that relatively simple changes in the continuous 

(s-domain) pilot model can produce large increases in the complexity of 

the corresponding discrete NIPIP estimation equation. To illustrate 

this situation it is useful to compare the two parameter model for the . series h, 8 model analyzed in Ref. 9, p. 61 with the above five parameter model. The two parameter model consists of a pure gain pitch 

attitude pilot element Yp and the same integral form of Yp. as used in 

8 h 

the five parameter model: 

TR-1l97-1 43 
(15a) 

(lSb) s 

The two parameter model may be z-transformed (using tables) directly 

into a two parameter estimation equation with a bias (Eq. 50 in Ref. 9) 

(16a) 

Furthermore, the estimation coefficients a and b may be directly and 

explicitly related to the unknown parameters in the continuous model: 

a -Ke (16b) 

(16c) 

For the five parameter model the situation becomes much more complex. The first complexity arises in treating the effective pilot time 

delay Tp It was found to be useful to represent the total time delay as 

Tp To + t.H (17) 

where 

TO kT 

and 

k 1, 2, 3, .... 

T data sampling time 

The incremental time delay tJ.T was then approximated as a first order 

lag, i.e.: 

e-tJ.TS tJ.T (18) (s + tJ.T) 

Estimates of the pilot time delay expected can be used to select a value 

for the integer k to make the value of /::':r as small as possible, thus 

TR-1l97-I 44 
improving the accuracy of the first order lag representation. The 

resulting pilot element is thus: 

( 1 /TL) e -1" OS 

Kpe (l/TI) (1/1:.1")" (19) 

The second and greater more difficult problem occurs in transforming 

the cont1.nuous s-domai.n model to first a z-domain representation and 

then to a difference equation. To make the algebraic manipulation 

involved tractable, the desire to maintain a direct relationship between 

the estimation coefficients and the unknown parameters in the original 

conUnuous model must be foregone and only the structure (Le., proper 

subscripts for 6, 6, and h) of the model retained. The estimation 

coefficients become undetermi.ned coefficients with no direct relati.onshi.p back to the continuous parameters. The esti.mation equation for the 

five parameter model (derived in Appendix A) 

. . . + cohn-k + q hn-1-k + c2hn-2-k -I- B (20) 

has nine unknown coeffi.cients. In spite of this relative complexi ty, 

the NIPIP algorithm applied to the STS-4 data converges and computes the 

estimation coefficients which appear valid at least on the basis of the 

correlation parameter. 

The problem at that point is one of interpretation of the identified 

model"s frequency response. It was found in the Phase III work that use 

of the w' transformation (Ref. 16) provided a good solution to this 

problem. After running NIPIP to obtain the coefficients, estimation 

equaUons in the z-domain may be transformed to the w' domain through 

= w'-2/T 

w'+2/T (21) 

where T-1 is the data sample rate in Hz. The value of this transformation derives from the fact the w' frequency response approximates that 

of the s-domain well below the sample rate, i.e., 

1'R-1l97-1 45 
G(w' ) G( jw) for w, , w , < 21f IT radl sec 

The overall procedure developed for representation and intepretation of 

complex pilot models in NIPIP may be summarized as: 

• Form continuous analytic model (s-domain) 

• Transform to z-domain 

• Simplify, preserving model structure 

• Transform to difference equation (estimation equation) 

• Run NIPIP and obtain estimation coefficients 

• Transform from z-domain wo w'-domain 

• Interpret w'domain response as jw frequency response for 

w « 21T IT 

An example of a NIPIP result as interpreted in the w' -domain is 

shown in Fig. 20. The difference between the s-domain and w' -domain 

frequency response is only a few percent at 10 radlsec and is even 

closer at lower frequencies. There is considerable uncertainty about 

the validity of this solution for the pilot model based on large changes 

in the model result as the fixed portion of the pilot time delay TO is 

varied -- a result not entirely unexpected due to the approximation 

of tn as a lag. However, further investigation of the behavior of this 

solution indicated certain data problems that must be resolved before 

further progress can be made in using NIPIP. 

D. SUMMARY OF FLIGHT DATA PROBLEMS 

AND RECOMMENDATIONS FOR OFQ 

The flight data analyses done in Phase III have revealed a number of 

shortcomings in the available data, Le., that in the (STS-4) MMLE file, 

TR-1l97-1 46 
20 

IYPl~ (Wi) I

dB: 

0.1.0 

..J 

1.DO 

I 

W I (rod/sec) :lLO.DO 

I 

o -'I.o--------------.. -----.m:...=-=:..:-::::..:-=.:-::::.f --

-20 _ 

o 

4- Y (w') Pe 

·-1.00 

(deg) 

·-200 

Figure 20. YPO From NIPIP Interpreted in w' Domain 

TR-1197-1 47 
which must be remedied for the OFQ. Fortunately, it appears that the 

needed data are probably available, the problem is primarily one of 

extracting the relevant variables from the large amount of data available from a number of sources. The primary problem identified in Phase 

III is the need for better altitude and sinkrate signals and it appears 

that this problem could be resolved through the use of available cinetheodolite data. This section will review the data problems, summarize 

OFQ data needs and present recommendations. A primary recommendation is 

that the Shuttle flight data be assembled into well-documented 'OFQ 

archive data ftles' containing all variables needed for vehicle dynamics 

and control and flying qualities studies. These files would be treated 

as an important OFQ ' product' and be available on the DFRF Cyber computer for use by local and outside groups. 

1. Altitude and Sinkrate Data Problems 

Altitude and sinkrate are of secondary importance for aerodynamic 

parameter identification, the primary use of the MMLE data files to 

date. However, they are essential for analysis of piloting technique in 

landing, either by phase plane methods' or the NIPIP program. In addition, the definition of effective body reference point becomes very 

important at altitudes on the order of the Shuttle body length. The 

altitude data is available on the MMLE file from two sources, the IMU 

channel and radar altimeter channel, and there is some inconsistency 

between the two. There is no sinkrate data available from the lMU channel and the radar sinkrate channel on the MMLE file was found to be 

unusable, due apparently to calibration problems, thus a computer file 

was created from Fig. 15 in Ref. 17. 

Figure 21 compares the radar and IMU altitude traces for the final 

30 sec of the landi.ng from the MMLE file. It may be seen that there are 

some differences particularly at low altitudes. To avoid the dynamic 

range problem i.nherent in looking at the altitude signal over the entire 

period from the start of pre flare Fig. 22 shows radar altitude for the 

shallow glide and final flare. The linearly interpolated signal indicates some ballooni.ng just before touchdown. A bias of several feet 

below the runway is i.ndicated at touchdown but this could easily be 

TR-1l97-1 48 
....:J 

:;Ci 

~ I 

I-' 

I-' 

\D 

-...J 

I 

I-' 

2000r 

I 

I 

1500 

-+-

........ 

~ 

Q) 

"0 1000 ~ :::J 

...... \D +-

« 

500 

5 

Pre flare Pullup -----~ .. ~I--- Shallow Glide -+- Final Flare ---l 

10 15 

Radar 

--- IMU 

20 

Time (sec) 

25 

Figure 21. Comparison of Altitude Time Histories from 

Radar and Hrn Channels in DFRF STS-4 MMLE Files 

Touchdown 

30 35 
Preflare 

Pullup --I-+-___ Shallow Glide 

90 

80 

70 

60 

,-

+-

-.-

-0:: 50 ~ 

~ 

Q) 

"'0 

:J 

..... 40 ..... 

<{ 

30 

20 

10 

Final Flare 

I Hz sample rate 

linear interpolation 

Touchdown 

OL-----------~-------------L----------~~~~~L-----L15 20 25 

Time (sec) 

30 35 

Figure 22. Interpolated Radar Altitude, Shallow Glide and Final Flare 

TR-1197-1 50 
accounted for by uncertainties 1n the body pos1t10n reference for the 

signal. 

The radar sinkrate plot created from the computer file generated 

from Ref. 17 is shown in Ftg. 23 (dashed line). This signal was generated at JWRF by di fferentJating the radar altimeter signal, and thus 

retai.n.s artifacts of the 1 Hz sample rate. For compar1son, a sinkrate 

trace was computed from 

. 

h "" VT sin y 

VT sin (8 - a) 

and is also shown (dotted line) in Fig. 23. When compared to the radar 

der1ved signal, there 1s cons1derable d1fference in basic dynamic characteristics as well as absolute levels over substantial periods. 

Given the inconsistency between the available data, a sinkrate signal was generated by complementary filterlng normal acceleration from 

the ACIP accelerometers and altitude from the lMU source. The filter 

form is shown in the block diagram of Fig. 24a. It may be seen that the 

complementary filtered sinkrate consists of washed.out altitude (Le., 

low frequency differentiation of altitude) combined with a pseudointegrated accelerometer signal as indicated in the Bode amplitude 

sketch of Fig. 24. The expression for the filtered h is 

(23) 

It may be seen that the high frequellcy accelerometer noise and low frequency altitude errors are filtered out. This s-d.omain expression was 

converted to a difference equation for use with the digital data, i.e., 

TR-1l9l-1 

. 

he n 

51 

-aT I-e· 

a· (24) 
15 20 

Time (sec) 

25 o ~----~----~~r-.------.-----~~--~~--~r----r.~--~~ 

-51----

-10 

Final Flare ----if I ~ : I I 

.... - 20 --- _: ___ 1 ------------1-- ·f~ / J I 

~ -25f-Ll- c : I I 

I 

~-~--·-t [ I 

l__ ___ _ _! ___ I -- - , . _-_ ... _- --..... - .... -

--- Radar Ii From Ref. 17 

•••••• ti = V sin (8-a) 

Goal 

T 

I 

---tTouchdown i 

! 

I 

! 

-----

i 

, 

I 

I 

: .... __ . +------

I 

, 

j 

I 

ri 

i 

I 

I L_ 

I 

I 

I 

i 

I if) ~ r I . -30 .: ------1 j---------I-------r- V,8,a From MMLE File , L 

I 

• I 

: I 

-35 ~ .. -------+--. --'---J 

-40 --~~·~~--·i-····· 

Complementary Filtered h : 

I I I , 

, 

I 

I I 

I I I --------r------ - -- i 

I 

L 

Figure 23. Comparison of Sinkrate Signals 

TR-1197-1 52 
no 

~LrD-~ 

nn +t------

~*"~ aJ Filter Block Diagram 

0=1 

I ------7-----

Altitude ~ '-.... -\(0 Acceleration 

Washout "'" Pseudo -I ntegrator ~ 

t 

~ 

~. 

~ 

Log~ 

bJ Frequency Domain View 

Figure 24. Sinkrate from Complementary Filter 

TR-1197 -1 53 
The complementary filtered sinkrate, compared to the other sources 

in Fig. 23, shows considerable smoothing with respect to the radar 

derived trace with more nearly constant sinkrate in the shallow glide 

region and lower sinkrate through the flare including negative sinkrate 

(Le., climb) beyond t == 28 sec. This indication of ballooning is qualitatively consistent, except for time of occurrence, with the radar 

altitude trace shown in Fig. 21. The complementary filtered sinkrate 

appears to lie generally between the radar and computed sinkrates. 

To emphasize the importance of high quality sinkrate data the phase 

plane trajectory using the radar sinkrate (dashed line Fig. 23) is shown 

in Fig. 25. An interpretation of this plot might allow one to consider 

it consistent with the basic pilot model and Fig. 18, however, it is 

much less obvious or convincing. 

2. Other Data Questions and Needs (STS-4 MMLE File) 

a. Attitude (Euler) Angle Frequency Response 

The attitude signal used in the NIPIP identification of the pilot 

pitch control element was the pitch attitude signal from the onboard 

GPC attitude processor. Pitch attitude should be consistent in frequency response with the integral of q from the ACIP (except for p and r 

contributions). Lateral directional motions are small in the shallow 

glide, however, there does appear to be some difference in frequency 

response in the pilots crossover region (Fig. 26), particularly in phase 

angle, which has not yet been explained. 

b. Earth Referenced X and Y Position 

In addition to altitude, the location of the Shuttle with respect to 

the Earth's surface (X, Y) is needed but not available in the (STS-4) 

MMLE file. X position is of particular interest for the landing to conveniently and precisely define landing distance for each flight. 

c. Manual Control 

RHC deflection data appears to contain a calibration uncertainty 

which may be due to the use of the backup flight control system as the 

TR-1l97-1 54 
u 

Q) 

V'l 

....... 

+-

'+-

• ..c 

, 

Q) 

+-

0 

.... 

...x 

c 

(f) 

.... 

0 

"'0 

0 

0::.: 

A I titude, hR (ft) 

o 10 20 30 40 50 60 70 80 90 O~-----~----~----~----~----~----~----~----~-----r5 

10 

15 

~~O 

:::5 

~,o 

Figure 25. Altitude-Sinkrate Trajectory. STS-4 Shallow Glide 

and Final Flare, 'h from Reference 17 

TR-1197-1 55 
(deg) 

o.-----------------------------------------------------~ 

-40 

-60 

-100 ._-

-200 

-300 

~ ........ Processor 

. £ Attitude 

o 

Integral of 

ACIP q 

.......... o "-

::::()-.. -.. ~ r Attitude 

"~_ Processor 

~ 

" , 

Integral of " 

ACIP q 

\ 

"-

\ 

'0 

\ 

\ 

o 

-400~------------------------------------------------~ 

0.1 

TR-1197-1 

1.0 

w (rod/sec) 

Figure 26. Frequency Response Comparison of 

8/0RHC from Attitude Processor and ACIP 

56 

10.0 
sourcE~ for thIs data. Further, the highest available 0RHC sample rate 

is 12.5 Hz (BFCS) rather than the 25 Hz actually used in the flight cant rol system and this signal is the sum of the pilot's and commander's 

inputs. 

There are no speedbrake or body flap control signals or manual trim 

data available in the STS-4 MMLE file. 

d. I"CS Mode Discretes 

There are some limited data available in the MMLE file for switching 

discretes. This data has not been exercised but appears to be adequate 

to define AUTO/MANUAL status for the pitch, roll, yaw, body flap, and 

speed brakes channels. However, there are a number of discretes of 

interest (Table 9, Ref. 2) which are not available. 

3. Summary of OFQ Data Requirements and Sources 

Table 4 summarizes the present view of the data needed for the OFQ 

and the available sources (indicated with X' s). The first four columns 

relate the data available on the DFRF MMLE files to onboard sources. 

The last four columns indicate data availability from four sources which 

have been investigated but not used to date. It is felt that these 

additional sources should be adequate to resolve most of the data problems noted above. 

a. Theodolite Measurements 

Problems with altitude and sinkrate data in, landing are of particular concern. It is proposed that available theodolite data be used to 

augment onboard altitude data. Theodolite data is available from Air 

Force Flight Center (AFFTC) on magnetic tape in digital form and it 

should be possible to enter this data into the DFRF Cyber computer with 

reasonable effort. Data is available from the Cant raves cinetheodolite 

system (Ref. 18) for 8TS-l through 6 and provides position, rate, attitude, speed, and wind data. Dati is also available for ST8-4 through 6 

from the 'takeoff and landing towers' parallel to EAFB runway used for 

TR-1l97-1 57 
TABLE 4. Su~Y OF SHUTTLE FLIGHT DATA DESIRED Al~D SOURCES FOR OFQ 

MMLE FILE GROUND BAS 1m THEDOLITE BEST ESTIMATED 

I MEASUREMENTS TKAJECTORY DATA NASA JSC !-!PDB 

(DFRF & LRC) 

VARIABLE ACIP GPC 01 BFCS TAKEOFF & LANDING TOWERS I CINETHEODOLITEi ! -.. -

t I 

-

r I i ALL DATA 

TRANSLATIONAL ACCELERATION X X I X ! EXCEPT I i 

AZ Ay AZ (NO AX) I \ \ I 

AeIP 

=F .. --------.. --_ .. 

ANGULAR ACCELERATION -----1------------1 I 

I 

i 

X I I 

P,Q,R I 

, I 

I 

--- t ' I i i I I ; 

i I I ! 

I 

TRANSLATIONAL RATE 

I (AIR .DATA) X I I X X \ 

a,~,H,VTRUE'VEAS'q,M i 

I I ---------------1 ------------t I 

I I 

TRANSLATIONAL RATE 

I 

I i 

I 

(E~TH. REFERENCED) X X , X 1 x, Y, Z 

I 

I I 

ANGULAR RATE P.Q.R X X I "--- ------- ---------i-----

X I 

EARTH REFERENCED POSITION X. Y. H, (Hrnu X X X 

I 

HR only) 

-.----

EULER ANGLES 1jJ, e. 4> 

i X X \ 

--- "----

I 

CONTROL SURFACE oe. ca. or X 

DEFLECTION I °SB X , 

°BF X i 

------- i .. --- -

MANUAL CONTROLS OqRHC' °PRHC Xl X2 

(COMMANDER AND OPED X I 

PILOT SEPARATE) °SBC. °BFC i 

I - ---.-f-------- ---- ------ ---------------f------------ -- ------- -- .. _--_._--- - ----- +----

i 

MANUAL TRIM CONTROLS ! --- ------ -- ----_._-- .------

SWITCHES AND FCS DISCRETES X 3 

----------- --_._-------- _._- ------------- I 

DISPLAY AND HUD VARIABLES 

----- ---- ----- r--- ---f------- --- ---------- --.-- .. _------- ---------"----- l 

WIND DATA X X + 

1 Pilot's input, 1 Hz sample rate 2 Pilot + commander input, 12.5 Hz sample rate 3 AUTO/MANUAL FCS status only 
the Shuttle. This data is somewhat more accurate than that from the 

Cine system, but is limited to X, Y, Z positions and rates. There are 

apparently no plans to obtain theodolite data for Shuttle landings at 

Cape Kennedy and it is proposed that the feasibility of obtaining such 

data be seriollsly considered. 

b. Best Estimated Trajectory Data 

Contacts with personnel at NASA LRC have been made concerning the 

Best Estimated Trajectory (BET) tapes and documentation has been 

requested but not yet received. In the interim, however, it has been 

learned thClt data already available at DFRF (Ref. 19) may be more suitable. This work merges data from four sources through use of a linearized Kalman filter to obtain a best estimate of vehicle and wind velocities. This effort is particularly useful because it is being done at 

DFRF, it sllOuld relate well to the MMLE files, and it is unusually well 

documented. It should be possible to extend to additional variables 

which could be useful in resolving data frequency response consistency 

problems (e.g., Fig. 26). 

c. NASA JSC Master Products Data Base 

A final source of data which was :i,nvestigated in the Phase III 

effort is the official post-flight data from the NASA JSC 'Master Products Data Base' (MPDB). The data flow and maintenance of this system is 

indicated in Fig. 27. Certain problems have made use of this data 

source impractical to date. First, the MPDB is extremely large and general and includes much data irrelevant to this program. Thus, adequate 

documentation is critical to extracting .the small subset of relevant 

data. A large number of specialized 'data products' (e.g., tapes, tabulations, plots, etc.) are generated from the MPDB after each flight for 

specific users and computer generated reports listing the specific variables available on each tape, etc., have been received from JSC, Ref. 

20. These reports (a typical page is shown in Fig. 28) are, however, 

not appropriate for specifying a request for a unique OFQ tape. What is 

needed is an organized listing of the variables available on the MPDB 

TR-1l97-1 59 
TAPE COMPARES 

(Flight-to-Flight) 

SHUTTLE PRODUCT 

DATA CONTRO REQUIREMENTS 

g BOARD I----~ - NEW PRODUCTS 

(SDCB) - UPDATES 

INTERFACE 

LOAD/CAL 

TAPE 

UPDATES DESIGN 

VERIF 

MISSION ANALYSTS 

PARAMETER 

FILE 

(MSID INFORMATION) 

I 

I 

1 

I 

L __ 

1 

I 

I 

I 

--_--1 

Figure 27. NASA JSC Master Product Data Base for Shuttle 

REQUEST 

CAL BLOCK 

UPDATES 

-500-VERIF 

-601-t~SFC 

SSD 

MASTER 

PRODUCTS 

DATA 

BASE 

(MPDB) 
'; ( ';(i (;(;1 REPORT AS or 15 MAr~ 19f1:-l PAGE 127 

MS ID NUMENCLATURE CALF LlrL UNITS 

CGf) lor LACY 

V9fl+L'0141'C 7--VELOCITY BFS Y FT/s 

V9fl'W'OI5()'C r;PEFNWICH MEAN TIME BFS Y US 

V9S+X'0431'X IACAN 1 RANGE DATA GOOD BFS Y EVENT 

V98+X'0437'X lACAN 2 RANGE DATA GOOD BFS Y EVENT 

V98'X'O!J33'X 1 ACAN 3 RANGE DATA GOOD BFS Y EVENT 

V9H'X'O,1J,-1+X TACAN 1 BEARING DATA GOOD BFS Y EVENT 

V9fl'X'OI13,,'X TACAN 2 BEARING DATA GOOD Brs y EVENT 

V9fl+X'O,1JG'X TACMJ 3 BEARING DATA GOOD BFS Y EVENT 

V9fl'W'14JO'C ACC SENSED VEL TIME WORD 1 BFS Y S 

V98'R'1550'C liOLl l~ATE RFS Y OEG/S 

V98'R'1551'C PITCH RATE BFS Y OEG/S 

V98'R'155:"C YAW RATE BFS Y DEG/s 

V98'H'15"O'C ROLL EULER ANGLE BFS Y OEG 

V9fl'H' 1561 'C PI TCH EULER ANGLE BFS Y OEG 

V98'U'I590'C ALPHA BFS Y OEG 

V98 'll' 15'3 I 'C BETA rlFS Y OEG 

V98'U'I597'C HEADING REL TO MAG NO BFS Y OEG 

V98'L'1790'C EARTH RELATIVE VELOCITY BFS Y FT/S 

V98'W" 18flR'C f I ME fO MAIN ENGINE CUTOFF BFS Y S 

V9S'L"19P'C MACH BFS Y ND 

V9fl'IF:? 160'C rMV DERIVED ALPHA BFS Y [)EG 

V9R'U'2340'C r-JAV DF.RIVED Q BAR(O) BFS Y LBr/FT2 

V98 '11">2350 'C 1101 ROLL SINE POSN BFS Y NO 

V9S*f-j '2351'C ADI PITCH SINE POSN BFS Y NO 

V98'f-j'235;>'C AOI YIIW SINE POSN BFS Y NO 

V9B'H'235~i'C AOI ROLL COS POSN Brs y ND 

V9B'H'235G'C AOI P ITCH COS POSN BFS Y NO 

V90'H'2J57'C IIOI YAW COS POSN BFS Y NO 

V90'U' 2,10[1' C MAJor, MODE BFS Y NO 

V9U'H'3100'C rlL SEL TACAN r~ANGE BFS y FT 

V9fl'H'3110'C rIt. SEL TACAN BEARING BFS Y RAD 

V9R'P'3117'C (l'(NAMIC PRESS flFS Y LBF IFT2 

V9fl'U'315r,'C ALPHA CMD BFS Y OEG 

V98'H'3160'C ROLL ANGLE CMO BFS Y OEG 

V98'U'314,)'C lACAN AZIMUTH PESTDUAL BFS Y OEG 

V'1B'lI'J,l,lG'C TACAN RAN(iE RESI~UAL RFS Y NM 

VC18'H'376;' 'C rllCAN 1 BEARING DATA ElFS Y OEG 

VQO '.1' 37GB 'C TlICAN 1 RlINGE A firs y NM 

V9O'II'J80~j'C lACAN 2 BEARING DATA BFS Y OEG 

V98'IP38S,l'C TACAN 2 RANGE A BFS Y NM 

V98'H'J<JOS'C TACAN 3 BEARING DATA BFS Y OEG 

V90 'Ht 395,1'C TACAN 3 RflNGE A Rrs Y NM 

V9U'W'395G'C lACAN TIME TlIG Brs Y S 

V99'U t 3,150'C CI. USTfR TO MSO/IMtJ-l COMPONENT 1 . 1 Y NO 

V99'll'3451'C ClUSTER TO M50/IMU-2 COMPONENT 1,1 Y ND 

V99'U'345;>'C CLUSTER TO M50/I MU - 3 COMPONENT 1.1 Y NO 

V99'U'34SJ'C CLUSTER TO M50/IMU-l COMPONENT 1.2 Y NO 

V99'U'3!J54'C CLUSTER TO M50/IMU-2 COMPONENT 1.2 Y NO 

Figure 28. Typical Page from the STS-6 Computer Compatible 

Tape (CCT) Report for the NASA JSC MPDB 

TR-1197 --I 61 

TYPE 

SPL 

BSU 

BD 

BO 

BD 

BD 

BD 

80 

EPL 

I1PL 

HPL 

HPL 

I-IPL 

HPL 

flrL 

IfPL 

flPL 

SPL 

SPL 

SPL 

HPL 

flPL 

HPL 

HPL 

HPL 

HPL 

HPL 

HPL 

HXU 

SPL 

SPL 

SPL 

HPL 

HPL 

SPL 

SPL 

RSLJ 

BCD 

BSU 

BCD 

BSU 

RCD 

OPL 

SPL 

SPL 

SPL 

SPL 

SPL 
with precise definitions and their 1'1SIO numbers. Further definition of 

dRta characteristics (e.g., sample-rates, sensor locations, corrections 

for time skews, etc.) is also necessary. It is believed that documentation of this sort exists and efforts are being made to obtain it. However, in addition to documentation difficulties, discussions with programmers at DFRF have indicated that the JSC tapes are written in UNIVAC 

formats difficult to read on the DFRF Cyber computer. Some data tapes 

presently are available at DFRF and may be read onto the Cyber; however, 

it appears that these tapes do not have much of the data required for 

the OFQ. 

4. Recommendations for Obtaining OFQ Data 

An initial task for the OFQ, which should be accomplished as soon as 

possible is the assembly of the required flight data from the above 

sources into a form that may be used with high efficiency. It is anticipated that 'OFQ archive' files can be created for each flight beginning with STS-l and made available on the DFRF Cyber computer. These 

files would be considered an important 'product' of the OFQ ~lich could 

be used at DFRF or (through transmission over phone lines) at remote 

facilities. This would avoid the all too common situation in which 

flight test data, obtained at great expense, is either destroyed or 

becomes effectively unusable shortly after the original study. 

It is proposed that these archive files be based on the DFRF MMLE 

files augmented with trajectory and wind data from the Ref. 19 work and 

the AFFTC theodolite data. Data from the JSC MPDB may be added as practical and as needed • The OFQ archive files can probably be limited to 

. the region from M = 3 to touchdown and rollout, at least initially. 

Figure 29 shows the envisioned generation and use of the archive files. 

The 'data merge' program must be developed, but it is believed that this 

could be an expansion of existing DFRF programs. It will be essential 

(but difficult) to have the important aspects of the OFQ files defined 

and documented (perhaps including basic documentation on the Cyber). 

This should include 

TR-1l97-1 62 
STS2·· __ 

STSI -

MMLE 

File 

(tape/disk) ~ 

DFRF 

Estimated 

Trajectory Tape 

(LKFSTSI, .... ) 

AFFTC 

Theodolite 

Tapes 

JSC 

MPDB 

CCT 

. . 

1--- ---------------------, 

I Data Merge Program , I 

I I 

I --, ---, I ----

I DF R F OFQ Archive. I 

CYBEli Files" I COMPUTER ______ ~ ____ ~ I L ___________________ ~ 

r-------------------~ 

I ~ta Selection a~ I 

I Transfer progra~ I 

I I - I 

I [Local Archive] I I File I 

I I 

I REMOTE I Local Working 

I COMPUTER Files I 

FACIL.ITY L----_______ ' _________ .-.1 

Figure 29. Generation and Use of the OFQ Archive Files 

TR-1l97 ·-1 63 
• defini.tion of each variable with name, code, and 

MsrD number when appropriate 

• origi.nal data source (e.g., ACIP, etc.) and/or 

post flight processed source (e.g., MPDB, etc.) 

• auxiliary information (sensor location, axis system, reference points) 

.. units 

• data characteristics (sample rate, resolution, 

accuracy) 

• applied processing (interpolation, filtering, 

time skew correction, bias corrections) 

Reference 19 is perhaps the best model of the level of documentation 

desired. The archive files should be created for all flights beginning 

with 8T8-1 and continuing through the end of the OFQ. 

TR-1l97-1 64 
SECTION III 

CREW QUALITATIVE ASSESSMENTS 

A key initial, and continuing, goal of this program has been to conduct in-depth, face-to-face debriefing interviews with the individual 

flight: crews. Ideally the debriefing would take place as early as possible after each flight to discuss the events, maneuvers, etc., examined 

on thl~ flight records, to solicit suggestions and recommendations concerning any additional flight segments requiring analysis, and to obtain 

commentary or pilot ratings pertaining to vehicle flying qualities and 

flight: control system performance. In practice, this has not yet been 

successfully accomplished for several reasons. One pertains to the many 

responsi bili ties, complex sc.hedules, and general NASA sequestering of 

the crews which make it difficult to gain access for in-depth interchangE!. Others pertain to the time lapse between the flight, the availability of flight traces suitable for flying qualities analysis, and the 

actual analysis of the flight records. 

It: has therefore been necessary, to date, to extract possible flying 

qualities information from crew comments volunteered in the general systems debriefing held at the Johnson Spacecraft Center (JSC) or in other 

press releases (e.g., Aviation Week and Space Technology). This information is then used to help interpret the flight traces (as in the previous secti.on) and to formulate a detailed questionnaire which will be 

used to guide the actual flying qualities debriefing when, or if, it 

should come to pass. 

The following subsection summarizes the key comments obtained from 

the STS-5 and -6 systems debriefings and associated press releases. 

Similar information from STS-l through -4 is contained in Ref. 2. The 

second subsection outlines considerations which led to the latest version of the questionnaire for crew in-depth debriefing. The revised 

questionnaire is contained in Appendix B. 

TR-l197-1 65 
A. STS-5 AND -6 CREW COMMENTARY 

1. STS-5 

This entry and landing was initially planned to be flown essentially 

all automatic (autoguidance and autoland). Per Ref. 21, very brief 

periods of manual control were scheduled during the steep approach 

(outer glides lope) segment to provide the crew with a "feel" of the aircraft in case they had to take over manually during the autoland segment. The guidance system incorporated the new Optional Terminal Area 

Targeting program which provided a "shrinking" Heading Alignment Cone 

(instead of cylinder) for improved energy control during the descending 

turn. A computer graphics generated "shrinking HAC" was displayed on a 

CRT to augment the flight director needle display. This flight was also 

the first to use the "ball/bar" Precision Approach Path Indicator (PAPI) 

on the ground to insure proper execution of preflare and final glide. 

The external visual view is that the "ball" will approach a bar of 

lights at a specific rate if the approach is being flown properly. The 

aircraft commander (Brand) views the external reference while the pilot 

(Overmyer) calls out altitude and airspeed. Thus the new aids were 

planned to assist the crew in their primary function of monitoring performance of the autoguidance and autoland systems. 

A few weeks before launch the above plan was revised, because of 

unresolved discrepancies in autoland performance between the mission 

simulator and shuttle training aircraft, to allow manual control from 

about Mach 1 on to touchdown. The landing was also rescheduled from the 

lake bed to the EAFB runway 22 due to moisture on the lakebed. 

In the STS-5 Systems Debriefing at the JSC, Brand indicated that he 

waited until about 0.85 M (and 40,000 ft) to take over manually due to a 

very pronounced shaking of the aircraft from Mach 1 to 0.85. Specific 

comments covering the manually controlled descent and terminal phase 

included: 

"at 30 to 40,000 ft the orbiter handling qualities 

are very pleasant" 

TR-1l97-1 66 
"Shuttle felt 'crisper' than the STA" 

"Easy to follow the error needles" (during the HAC; 

275 Kt and 1.6 g) 

"The handling of the orbiter during pref lare and 

flare felt like the STA ••. The STA is a good trainer 

for manual landings" 

"It is good to fly the airplane from Mach 1 to landing as this gives continuous experience with the aircraft so that the pilot is comfortable flying it by 

preflare" 

" ••• would hate to take over at 300 ft" 

Brand indicated he mostly looked out the window after they broke 

through the cloud deck (on steep glideslope) ••• that the ball-bar indicator is a good aid for doing the preflare... that he had no problem 

knowing how much input to put in at the preflare (which he contrasted to 

Mattingly's uncertainty on STS-4 where the ball-bar was not available). 

The landing gear was lowered at 400 ft altitude instead of the origi~ally prescribed 200 ft (presumedly to minimize workload and/or perturbation during the final flare). Brand indicated he paid no attention to 

the roll needle while on final -- (Overmyer) called out airspeed and 

altitu.de -- (Brand) thought they were still 3-4 ft above the runway when 

the main gear touched down. 

In summary, the general impression obtained is that attitude control 

and flight director tracking is precise and relatively easy in the 

Shuttle. However, as the vehicle nears the surface and the crew becomes 

concerned with precise flight path control, the task becomes quite 

demanding, external visual aids are high desirable, and any path corrections, disturbances, or configuration changes during the last few seconds of final glide and flare can be troublesome. 

2. STS-6 

This was the firs t flight of Challenger, the second vehicle. It 

incorporated a Head-Up-Display (HUD) with improved visual information 

for the terminal phases of flight. This was flown manually from the 

TR-1l97-1 67 
Heading Alignment Cone (HAC) through touchdown (Ref. 22). The landing 

was again made on the EAFB runway 22 due to the lakebed being covered 

with water. 

The pilot (Bobko) commented in the systems debriefing that it had 

been their experience in the STA that if the orbiter got off the nominal 

trajectory in the HAC turn it was impossible to get back on nominal 

track using the HUD alone. Thus it was their strategy to use the HUD as 

another piece of information, not as a sole guidance source. The commander (Weitz) said that he flew the (head down) flight director error 

needles rather than the HUD. 

The HUD is also useful on the steep glide slope along with the 

flight director needles and out-the-window visual aids. White triangular markings that correspond to steep glide slope intercepts are 

placed on the ground (Ref. 14). According to the report of the debriefing the crew was told to use the high-wind, close-in aim point. This 

caused some confusion in that the crew was not sure whether or not the 

HUD guidance was displaying the high wind or nominal aim point. Weitz 

said that the HUD alm point was not where he expected it to be, and 

Bobko said it is not an easy task to decide where the aim point should 

be if the marker on the ground is not visible. (The lakebed aim point 

marker was washed out or covered with water.) Weitz found that the 

relationship between the HUD aim point and the PAPI light location was 

not as he was eKpecting, but by preflare the aim point looked good. He 

also said the HLJD helps a lot from 2,500 ft on down. The HUD displays 

among other things (Ref. 23) pitch and roll attitude, reference flight 

path, velocity vector, altitude, airspeed, cross track error, and speed 

brake position. 

The commander (Weitz) said that he did not fly the HUD tightly, that 

is he dld not tighten his gains to the point of putting the velocity 

vector on the guidance diamond and used the "normal" inside cockpit scan 

(Ref. 22). 

The crew said that the orbiter flew through gusts at 4,000 ft and 

again at 400 ft, and that it was unresponsive to the gusts. Weitz said 

TR-1l97-1 68 
that no pilot inputs were required as a result of these gusts. Overall, 

Weitz said that the "subsonic. handling qualities were more crisp than he 

was expecting... there was no sloppiness or overcontrol tendencies ••• 

the orbiter was a good, solid, nice flying machine ••• he had no recommendations for manual flight control improvements." 

B. R1WISED CREW INTERVIEW/DEBRIEFING QUESTIONNAIRE 

The format of the initial questionnaire (Refs. 1 and 2) were devoted 

solely to vehicle flying quality considerations. As our understanding 

of the uniqueness of superaugmented vehicle dynamics has increased and 

with the additional experience gained on each Shuttle flight has concluded, it has become increasingly apparent that 

a) each flight (landing) has been unique in terms of 

manual vs. auto control of various segments of 

the approach and landing, 

ground and airborne visual aids available to the 

pilot, and therefore 

closed loop strategy and techniques available to 

and employed by the crew, and 

pilot training, background, and experience in 

flying qualities evaluation 

b) we often do not know how to intepret crew comments 

and/or what we see in the flight traces 

Therefore the crew interview/debriefing questionnaire needed to be 

rearranged and expanded into two sections (see Appendix B). The first 

covers crew perception of the control tasks and flying techniques in 

each entry flight segment. This includes identification of closed loop 

structure employed, strategy in transitioning from one flight segment to 

anoth,er or from one closed loop structure to another, tasks which were 

perfo~med on a precognition basis (i.e., highly trained, open loop 

reaction), etc. The second section covers crew evaluation of the 

Shuttle dynamic response, workload, and flying qualities associated with 

the perceived control tasks. In both sections five flight segments are 

addressed: 

TR-1197-1 69 
Supersonic bank reversals 

Heading alignment circle or cone (HAC) turn 

Descent on the steep glide slope 

Preflare 

Shallow glide and final flare to touchdown 

The information obtained from the first section will be directly 

applicable in modeling the pilot/vehicle control loop structure and 

extraction (via the identification techniques previously discussed in 

Section II) of model parameters, pilot adapt ion (or settling) time, etc. 

The information from the second section should provide qualitative and 

quantitative data, together with data extracted from flight trace analysis, from which to establish flying quality and/or flight control criteria and design guides for future Shuttle craft. 

TR-1l97-1 70 
SECTION IV 

HAND CONTROLLER EXPERIMENT 

It was noted in Ref. 1 that the Shuttle Rotational Hand Controller 

(RHC) force/displacement characteristics differed drastically from those 

considered to be acceptable-to-good in past in-flight simulations 

(Refs. 24 and 25). The Shuttle RHC exhibits much greater displacement 

per unit applied torque than the largest ratio tested in Ref. 24 while 

the motion response command per unit torque was much lower. Concern was 

expressed in Ref. 1 that these characteristics might contribute to 

degraded flying quality ratings and also might lead to a larger effective time delay (latency) within the pilot's neuromuscular system. [It 

was shown in Ref. 26 that loose vs. stiff manipulator force/displacement 

characteristics have a significant influence on the human operator 

effective time delay in a tracking task.] 

Later it became apparent that the RHC force/displacement characteristics presented in Ref. 1 represented the ALT vehicle. Information 

contained in Ref. 27 indicated these RHC pitch and roll force gradiE~nts 

were considered undesirable, a possible contributor to the PIO experienced in the ALT-5 landing, and were doubled for the OFT vehicles. However, it may be noted from Fig. 30 that there is still a very large discrepancy between the force/displacement and response/force ratios of the 

OFT and those considered acceptable in the Ref. 24 flight tests. 

A simple experiment was proposed in Ref. 2 to quantify any difference in pilot time delay induced by the OFT RaC and a manipulator exhibiting supposedly "good" characteristics. It is based upon the STI 

developed Critical Instability Tracking Task (CITT) in which the operator attempts to stabilize a controlled element having a steadily 

increasing divergence. The experiment was further detailed as a part of 

this Phase III effort and preliminary runs were accomplished at the DFRF 

by DFRF personnel. The following subsection briefly describes the theory hehind the measurements, the experimental plan, and results which 

TR-1l97-1 71 
.!: 

I 

..Q 

....... 

U 

Q) 

CJ) 

....... 

01 

Q) 

3 

"0 2 

c:. 

o 

(!) 

Q) 

CJ) 

c:. 

o 

c.. 

CJ) 

Q) 

0:: 

Q) 

c.. 

o 

en 

.S! Fixed Small Large 

~ 

OFT 

L 

M 

H 

VH 

~ 

ALT 

Roll 

Pitch 

- ~ c:. <-------- L-______ ~ ______ L_~ ______ ~I_~ ________ _LI ________ ~I ________ ~I __ 

.2 ;4 .6 .8 1.0 1.2 

Xs/Ts (deg/lb-in.) 

Fi6Ul-e 30. Shuttle Orbiter RHC Feel/Command 

Gradient Characteristics i.n Landing 

lead to the conclusion that the Shuttle RHC configuration and force/ 

displacement characteristics are such that they increase the pilot's 

neuromuscular time delay. 

A. KEY CONCEPTS FOR ASSESSING SENSITIVITY TO TASK VARIABLES 

Operational manual control systems are typically designed to require 

far less than the pilot's ordinary limiting capabilities. Therefore 

performance decrements due to task variables such as manipulator or display dynamics are seldom observed except for extreme values of the variables. However, skill-factor decrements due to task variables do become 

apparent when the pilot is near his limiting performance. Thus intrinslc skill limits can be measured only under high task-induced stress 

condit ions which push the pilot to his limits. 

Particular control tasks can be designed to emphasize particular 

skill factors. The Critical Instability Tracking Task is specifically 

designed to measure the lag induced into a closed loop control task by 

TR-1l97-1 72 
the pilot. The task is shown in block diagram form in Fig. 31a (from 

Ref. 28). The output of a first order divergent controlled element is 

displayed to the operator as an error. The operator attempts to stabilize the divergence through application of a force (and/or displacement) 

to a manipulator. The divergence inverse time constant, A, is increased 

as a function of time and error magnitude until the operator can no 

longer maintain control. The A value at loss of control is the critical 

instability, A

c

' which is approximately the inverse of the dynamic 

delay, Te. Typical time traces of the rate of A increase and the various motion quantities are shown in Fig. 31b. 

The reason for the difference in pilot lag contribution with a stiff 

(force) vs. a free (position) manipulator may be observed from the 

detail,ed block diagram of operator dynamic elements shown in Fig. 32 

(from Ref. 29). On the right side of Fig. 32 it may be observed that 

the neuromuscular actuation system has two feedbacks. One is via the 

force sensing spindle/tendon organ ensembles directly to the spinal cord 

where the error signal in generated to further control the muscle. The 

second feedback path is via the proprioceptive (joint) receptors which 

sense the various joint angles. This information is fed back to the 

central system (brain) for integration and equalization and generation 

of a new motorneuron command which then progresses back down the spinal 

cord as a command to the spi.ndle/tendon ensemble. Thus the neuromuscular actuation system consists of an inner, force, servo system and 

an outer, position, servo system. 

delay than the outer loop. 

The inner loop inherently has less 

Typical differences in critical instability with different types of 

manipulator restraint and controlled element complexity are shown in 

Fig. 33 (from Ref. 30). The upper plot is for a controlled element 

cons is ting of a divergent first order lag. The middle plot is for a 

second order controlled element consisting of au integration and first 

divergl?nce. The bottom data point is for a third order controlled element consisting of a double integration and first order divergence. It 

can be seen from Fig. 33 that the critical instability (pilot lag) is 

influenced most by task order and secondarily by manipulator characteristics. However, manipulator characteristics become more significant as 

TR-1l9l-1 73 
--. 

i=O 1-

I 

utopacer I Increasing 

Instability 

AU) Control 

Error Human Force Divergent Output 

e Operator F Controlled 

Element 

-KA 

--S-A 

Concept: 

A(t) --Ac , th'e "Critical Instability" 

Interpretation: 

Critical Instability~ ------

Dynamic Delay 

m 

F, 

N 

e. 

em 

m, 

em 

rod 

sec 

o 10 20 

a) Block Diagram b) Typical Trial 

Figure 31. The "Critical-Instabilitylt Task (First Order) 

Output 

30 40 sec 
SJ~lem System 

Ir.pul E ftor 

I e 

!_ SErISORY MECHANlsr,:s 1 _______ CECITRAl ElE/.1ErlTS _______ +-__ NEUROt.1USCULAH t,CTU.'.TIOIJ __ ~--I 

I ~ SYSTEM 

Pfupfl.).::n .. '!I-Je PrCpnc,c( pflve 

Ink.;'Jlion {JOint} 

a/.J RC;:C~h .. fS 

Equ;';hlollon I .. xp e-rp,; I 

Integrol 

-I -~j- ,-

.. Ti's f' 1 

1

0 

r-- ~~ Proportional --j PO"JIO,eGi f' -- ~ Mular '-5;i:7" Muscle Mcrllpula!or 

1-

K' VI$IJol I \Jtti.~CfS Ct:nlro! P a Moler neuron led I D,n(1:'TII(s 

Sllm...:11 Ol'tles hoci'SSlrlg COfTl11l011d I or I COlr,mend -rMs r,1or.JPulctor 

D,sploy '- ar~d ol,d f- InlCC]fJl 1v e j, < ..L!. KMe 

e, Mech~Jn!sms ac I I I,I( 5 2 21;,,5 II e 

+ €, Relino 

r~ Foveal I !nll'C]ra1Ion I I r ,." 

, I 

--'"-[ ----- IT 5,.JIj -- .. \ t--~- .. j 

LJ I t.~i.:ChJ;II<;m i Role i I t~ WI,II L,;", Palh .... ·oy j le-lrc+TnllljSI r Molorneuron " , I Eyeb,,11 K'R se- rRs Commond '-_ "--oJ 

POSdl0n 

A, Oculomolol 

t.1uscles a, Oculomotor Spindle I Teenon 

- Qnj Eye - y CClmfTlund AcceleroliGn Organ Enserr,bies 

D,.nGrnl~~ System K~s2e-TAS ~ 14-

Kspls+Z,pkrcps 

Visual Cr.onllct (s + r;p) 

Equaillalion , I~J1"OgmUS ! 

Crcscfcedc Yb 

"-

Vestlbulor ond Klne~th(:IIC Sen'.ors Ve:.llt;l;!or 

- (J,l{j K (--~- + I) L Incor Motion Angular t .. ~oIJon KU1t'S!!1I:tJc fJ Psp ; K,JTr.~AS 

lKw(~T)VWs e-T1 S I(tlcqrut,on -Tr~)2-2~0j ar.d (T pi) --,- + --- +1 

--- ____ L- NI '~r~ W N 

O,G s' I iT, s + 1110,1 s+l) Lquuh..'ol1on 

l.1r.e;]( AncJuiur 

AcceJC(utlon VelocIties 

. 

Figure 32. Structural-Isomorphic Model of Pilot-Aircraft System 

Controlled 

Elerrenl 

Y, 

5ys.'e m 

O;.;t,:'t..:1 

-r;;-

I 
....... 

u 

(l) 

en 

"-

"'0 

0 

"-

u 

...< 

>-

-+-

..Q 

0 

-

en 

c 

0 

u 

-'-

t) 

7 KA 

S-A -- Range 

6 

. -- Median 

1 N = 5 Runs 

5 

KA 4 S(S-A) Ss = WRA 

O"i = 0 

3 X = NASA CR674 

Data 

2 KA 

s2(s-A) 

OL-----~----~------------------._----

TR-1l97-1 

Stiff 

Stick 

Spring 

Stick 

Type of Stick 

Free 

Stick 

Figure 33. Critical Instability VS o Type of Control Stick 

76 
the task order increases. The ~c decreases approximately 1 rad/sec 

(Me:' 0.025 sec) between the stiff and free stick configurations for a 

first order critical task. With the second order critical task the ~c 

decreases about 1.8 rad/sec (6Te ~ 0.205 sec). 

B. EXPERIMENTAL PLAN 

The initial intent of the experiment was to compare the Shuttle RHC 

with a sides tick configuration given good (HQR = 3) ratings in the 

Ref. 24 flight test. The only stick available for comparison was from 

the DFRF fixed base simulation cockpit and, unfortunately, the gradient 

adjustment range was insufficient to achieve the desired values. The 

maximum gradient achievable remained quite close to that of the Shuttle 

RHC. Therefore, the tactic adopted was to adjust the DFRF stick break- , 

out and gradient to "best" values for this experi.mental task as judged 

by DFRF research pilots. These values were somewhat different from the 

Shuttle RHC and did prov1de for some comparison. 

Possibly of greater s1gnificance, the geometric characteristics of 

the two manipulators were quite different. The Shuttle RHC configuration 1e approximately that shown in Fig. 34. The pitch axis pivot point 

is at the middle of the hand palm. The roll axis pivot is about five 

inches below the pitch pivot. Thus deflection of the RHC is accomplished by wrist movement. Stop to stop deflection (exclusive of an 

override) is ±19.5 deg in each axis. The DFRF simulation stick is 

fairly conventional with the pitch and roll pivot poi.nts roughly a foot 

below the grip. The stop to stop deflection was ±2 in. in both axes, 

therefore the hand grip displacement was considered to be mostly longitudinal or lateral translation and involved arm, rather than wrist, 

motion. 

In both cases the manipulator grip was centered in front of the 

pilot. 

A block diagram of the experimental setup is shown in Fig. 35. The 

Shuttle, RHC electrical output. was processed through its basic digital 

flight control system computation and shaping elements and extracted 

TR-1l97-1 77 
~ DISPLAY 

Neutrat 

pOlition 

Pitch motions 

Left --------

Rear 

Roll mote;"",:' 

t-' 

Neutral 

position 

Left 

Rear 

Figure 34. RHC Configuration 

r---+- Yp ~ MANIPULATOR 

GAIN 

cn 

J\DJ 

COMPUTER 

<DELAY> 

Figure 35. Manipulator Experiment Set-Up 

TR-1197-1 78 

r--
from the computer as a pitc:h or roll rate command. This digital computation introduced an effective time delay of 0.046 sec. To keep the 

experlment focussed on manipulator characteristics, a similar time delay 

was inserted between the DFRF simulation stick output and the input to 

the critical task test unit. A gain adjustment was placed between the 

computer output and the CTT for proper electrical scaling. 

The tracking task was single axis, e.g., pitch or roll independently, using the first order instability: 

KA 

s - A 

Five test subjects were employed. After initial task familiarization 

and training, each was given five groups of three trials with each 

manipulator. The critical instability, AC' scores were recorded at the 

conclusion of each trial and the subjects were challenged to make the 

highest possible scores. 

c. Rl~SULTS 

The experiment was accomplished on a non-interference basis with the 

subjects normal workload and the simulator schedule. Therefore the 

training sessions for some configurations were somewhat shorter than 

desired and learning effects were noted in 5 of 20 sessions as a general 

increase in AC scores through the 15 trials. 

overall results is considered to be minor. 

However, the impact on 

Mean and 10' AC scores for the pitch axis task are shown in Fig. 36. 

The L above a score identifies sessions where learning was evidenced 

throughout the 15 tasks. Figure 36 shows Lt· of the 5 subjects had better 

scores with the conventional stick but this could have ended as 3 of 5 

if subject 2 had more training with the Shuttle RHC. 

Mean and 10' AC scores for the roll axis are presented in Fig. 37. 

Again 4 of 5 show better scores for the DFRF stick but this could have 

been 3 of 5 if learning had levelled off. 

It might be observed that the scores obtained in this experiment 

are lower than those of Fig. 33. This is because of the 0.046 sec 

TR-1l97-1 79 
6 • Shuttle RHC 

* 

X DFRF Stick 

L 

* L* , flCT ! MEAN 

4 ! +t r leT 

Ac 

2 PITCH AXIS 

n= 15 

o~--~--------~------~--------~--------~ 

2 3 4 5 

Subject 

Figure 36. Pitch Axis GGT Raw Scores 

6 • Shuttle RHC 

Lt 

X DFRF Stick 

L ,* L 

ft 

, feT 

4 

f~ ! MEAN 

Ac ICT 

2 ROLL AXIS 

n= 15 

o~--~--------~--------~--------~--------~ 

2 345 

Subject 

Figure 37. Roll Axis GeT Raw Scores 

TR-1197-1 80 
complltati.onal time delay which became a part of the controlled element 

dynamics. The effectilTe controlled element dynamics seen by each test 

subject was 

K).e-0.046s 

- (s - ).) 

This could be expected to place the scores somewhere between the first 

and second order critical tasks of Fig. 33. 

The, effective pilot latency with the computational delay, 'd' 

removed is 

1 

'p >'c· - 'd 

The average values calculated across the five subjects (Figs. 36 and 37) 

are presented in Fig. 38. These show the Shuttle RHC resulted in 

increased pilot latency (delay) of about 25 millisec in pitch and 

20 millisec 1n roll. This may be due to the RHC force/displacement 

characteristics, the wrist type motion, or both. 

With the computat1onal t1me delay removed as above, the resulting.,p 

values can be compared with the inverse of first order critical instability task scores from Fig. 34. This is done in Fig. 39 and shows the 

DFRF simulation stick to result in pilot neuromuscular lags comparable 

to the unrestra1ned (free) stick configuration of Ref. 26. 

Shuttle RHC fares somewhat worse. 

D. CONCLUSIONS 

Aga1n the 

It is concluded on the basis of this prelimi.nary .experiment that the 

Shuttle RHC may be inducing extra lag in the pilot because of its force/ 

displaeement characteristics, wrist action pivot point location, or 

both. The difference in this exper1ment is about 0.020-0.025 sec more 

delay 1.n each axis with the Shuttle RHC. 

ThE! experiment has demonstrated that the simple CTT approach can 

measure the influence and could be employed as one task in optimizing 

manipulator characteristics for future shuttlecraft. 

81 
Tp 

.3 

.2 

.1 

6.:: .025 sec 

~:: .020 sec 

OL--\.---'----'------...1--...1--~-

S 0 S 0 

Pitch 

T = p 

Roll 

Pitch 

Roll 

Figure 38. Effective Pilot Latency; Shuttle vs. DFRF Simulation Controller 

TR-1197-1 

.3 

.1 

Pitch Roll 

Q) 

Q) 

.... 

LL 

Figure 39. Comparison of Experimental Results with 

Free and Stiff Manipulator Data of Reference 30 

82 
SECTION V 

RESULTS, CON~~USIONS, AND RECOMMENDATIONS 

A. MAJOR RESULTS 

o Shuttle flight data (STS-4) has been transferred over 

telephone lines from the DFRF Cyber computer to the 

STI PDP-II computer for analysis. This procedure has 

proven practical and efficient. 

• The superaugmentation model has been used to study 

the identification of the effective augmented pitch 

response of the Shuttle and explain results obtained 

from (STS-4) flight data. 

• Identification of the effective Shuttle pitch 

response in landing from STS-4 flight data has been 

performed using a spectral method (the FREDA program) 

and also by use of the NIPIP program. 

o The pilot-vehicle-task model for shallow glide and 

final flare developed in Phase II has been refined 

based in part on flight data analysis. 

o The altitude-sinkrate phase plane method has been 

further developed and applied to the STS-4 flight 

data. It has been possible to extract pilot-vehicletask parameters, in particular, Yo and Tffor STS-4. 

• An attempt has been made to use the NIPIP program for 

pilot model identification but satisfactory results 

have not been obtained. This is due· in part to 

procedural complications and data uncertainties for 

this application, but progress has been made in 

developing modified procedures (e.g., use of the 

w' transformation). 

• A number of data problems and needs have been uncovered in the flight data analysis effort which need to 

be resolved for the OFQ. The most critical of these 

is the need for better altitude and sinkrate data in 

landing. 

• Seve ral addi t ional sources of da ta (beyond the MMLE 

files) have been investigated and appear to be applicable to resolve the data problems. These include 

the AFFTC cinetheodolite tapes, the Kalman filtered 

BET data from DFRF, and the JSC Master Products Data 

TR-1l97-I 83 
Base. Additional work is needed to access these 

sources. 

B. CONCLOSIONS 

• The effective augmented vehicle identification effort 

provides support for the superaugmentation model and, 

in particular, indicates that the effective attitude 

lead is (l/Tq) rather than (l/T8Z). The approximate 

time-invariance of the q/qc response in the preflare 

through touchdown region has been verified. 

• Normal pilot RHC activity in landing apears to be 

adequate for identification purposes. This indicates 

the feasbility of the indirect non-intrusive OFQ 

approach proposed in Phase II. 

• The NIPIP program appears to be capable of vehicle 

identification from flight data. 

• Examination of the STS-4 time histories (in particularly the RHC trace) indicates (expected) variations 

in pilot technique among landing flight segments, 

discrete RHC inputs to transition between segments 

and pulsive RHC inputs in shallow glide and final 

flare. 

• The form of the STS-4 altitude sinkrate phase plane 

trajectory is consistent with the Phase II shallow 

glide and flare pilot-vehicle-task model. The initial conditions, ho and hf, are lower than the 

'nominal' values estimated in Phase II, but when 

STS-4 initial conditions are inserted in the model 

the results are consistent with the Yo and Tf values 

extracted from the phase plane analysis. The oscillatory nature of the h-h trajectory indicates pilot 

difficulty in accomplishing the landing. 

@ Procedures for pilot model identification with NIPIP 

needs further refinement and h-h phase plane analysis 

should be used for guidance in NIPIP model formulation. 

@ The CTT experiment has demonstrated that this simple 

simulation approach can measure the influence of different manipulators and could be employed as one task 

in optimizing manipulator characteristics for future 

Shuttlecraft. 

• The Shut tIe RHC may be inducing extra lag in the 

pilot because of its force/displacement characteristics, wrist action pivot point location, or both. 

TR-1197-1 84 
The difference in this experiment is about 0.020-

0.025 sec more delay in each axis with the Shuttle 

RHC. 

C. RECOMMENDATIONS FOR OFQ 

• First priority should be given to bringing the 

required flight data sources together into documented 

OFQ archive files on the DFRF Cyber. Before any data 

is formally analyzed, data checks for consistency 

(e.g., altitude from radar, IMU, barometric and theodolite sources), biases, calibration, etc., should be 

made with adjustments, filtering, etc., made as 

required. 

• Vehicle identificati.on should be performed for all 

flights to develop an ensemble average for q/oRHC. 

Both FREDA and NIPIP should be used to gain further 

experience with NIPIP in flight data applications. 

• Altitude/sinkrate trajectories should be generated 

and analyzed for each flight with emphasis on pilot 

technique variations among flights. The pilot model 

should be refined as required. 

• The pilot-vehicle-task model for landing should be 

implemented as a simple digital simulation (STI 

PDP-ll or Apple). A primary use of this simulation 

would be to analyze NIPIP pilot identification problems and procedures using a known pilot model. 

• When preliminary work has been done, pilot identification efforts with NIPIP should be continued for 

each flight. 

TR-1l97-1 85 
This Page Int.~ntionally Left Blank 
REFERENCES 

1. Myers, Thomas T., Donald E. Johnston, 

Shuttle Flying Qualities and Flight 

Study, NASA CR-170391, Dec. 1981. 

and Duane McRuer, iPace 

Control System Assessment 

2. Myers, T. T., D. E. Johnston, and D. T. McRuer, SI?ace Shuttle Flying Qualities and ~~ight Control System Assessment Study, Phase 

~, Systems Technology, Inc., TR-1187-1, Dec. 1982. 

3. McRuer, o. T. and E. S. Krendel, Mathematical Models of Human Pilot 

Behavior, AGARDograph No. 188, Jan. 1974. 

4. Hess, Ronald A., "A Rationale for Human Operator Pulsive Control 

Behavior," J. Guidance and Control, Vol. 2, No.3, May-June 

1979, pp. 221-227. 

5. Systems Technology, Inc., "Flying Qualities and Control System 

Alternatives for Superaugmented Aircraft," NAS2-1l388. 

6. Aerodynamic Design Data Book, Orbiter Vehicle, Rockwell International, Report SD72-SH-0060, Vol. 1, Nov. 1980. 

7. Hoey, Robert G., et al., AFFTC Evaluation of the Space Shuttle 

Orbiter and Carrier Aircraft -- NASA Approach and Landing Iest, 

AFFTC-TR-78-14, May 1978. 

8. Magdaleno, R. E., The STI Frequency Domain Analysis Routine 

(FREDA), Systems Technology~nc., WP-407-2, June 1978. 

':). Hanson, Gregory D. and Wayne F. Jewell, Non-Intrusive Parameter 

Identification Procedure User's Guide, NASA CR-170398, Apr. 

1983. 

10. Bendot, Julius S. and Allen G. Piersol, Engineering Applications of 

Correlation and Spe£-tral Analysis, John Wiley & Sons, NY, 1980. 

11. Mooij, H. A., W. P. de Boer, and M. F. C. Van Gool, Determination 

of Low-Speed Longi tudinal Maneuvering Criteria for Transport 

Aircraft with Advanced Flight Control Systems, National Aerospace Lab., NLR TR 79127 U, Dec. 1979. 

12. Requirements/Definition Document, Flight Control, Part 1 Configuration Performance and Functional Requirements, Report SD72-SH0105, Vol. 1, Book 2, Part lA, July 1977. 

13. Heffley, Robert K., Gregory D. Hanson, Wayne F. Jewell, and Warren 

F. Clement, Analysis of Pilot Control Strategy, NASA CR-170399, 

Apr. 1983. 

TR-1l97-1 R-1 
14. Tsikalas, G. M., "Space Shuttle Autoland Design," AIAA Paper 82-

1604-CP, Aug. 1982. 

15. Stapleford, Robert I.., Irving L. Ashkenas, Graham Dunstan, John J. 

Best, and J. Alfred Tennant, Analysis of Several Handling Quality Topics Pertinent to Advanced Manned Aircraft, AFFDL TR-67-

2, June 1967. 

16. Whitbeck, R. F. and L. G. Hofmann, "Digital Control Law Synthesis 

i.n the w' Domain," J. Guidance & Control, Vol. 1, No.5, 

Sept-Oct. 1978, pp. 319-326. 

17. Preliminary ~nalysis of STS-4 Entry Flight Data NASA TI1-81375, 

Sept. 1982. 

18. "Final Analysis Output Progam (FALOUT)," Kentron International, 

Inc., Memo. No. K4E-007B-81, 10 Dec. 1980. 

19. Whitmore, Stephen A., "Reconstruction of the Shuttle Reentry Airdata Parameters Using a Linearized Kalman Filter," NASA DFRF 

proposed Technical Paper, 1983. 

20. "Master Products Data Base," NASA Johnson Space Center DDC Accession No. Z06-5-000515, 21 Mar. 1983. 

21. Covault, Craig, "Liftoff Time Pivotal to Shuttle Events," Aviation 

Week and Space Technology, Vol. 117, No.8, Nov. 1, 1982, 

pp. 57, 59, 61-63. 

22. Covault, Craig, "Shuttle 6 Reentry Aids Program Data Base," Aviation Week and Space Technology, Vol. 118, No. 16, 18 Apr. 1983, 

pp. 18-20. 

23. Looney, B. J., "Pos t-DETAC HUD Monitor Description," Sperry Autoland Memo 57, 4 Nov. 1980. 

24. Hall, G. Warren and Rogers E. Smith, Fltght Investigation of 

Fighter_ Side-Stick Force-Deflection Characteristics, AFFDL-TR75-39, May 1975. 

25. Black, G. Thomas and David J. Moorhouse, Flying Qualities Design 

Requirements for Sidestick Controllers, AFFDL-TR-79-3126, Oct. 

1979. 

26. t1cDonnell, J. D. and H. R. Jex, A "Critical" Tracking Task for t1anMachine Research Related to the Operator's Effective Delay 

Time. Part II: Experimental Effects of System Input Spectra, 

Control Stick Stiffness, and Controlled Element Order, NASA 

CR-674, Jan. 1967. 

TR-1l97-1 R-2 
27. Hartsfleld, Henry W., "Space Shuttle Orbital Flight Testing," Society of Experimental Test Pilots 22nd Symposium Proc., SETP 

Technical Review, Vol. 14, No.2, 1979, pp. 11-20. 

28. Sadoff, M. E., D. T. McRuer, and H. R. Jex, "Human Transfer Function and Dynamic Response Properties During Extended Orbital 

Flight," Systems Technology, Inc., WP-175-19A, presented at 

NASA-Ames/National Aeademy of Sciences Experiment Jury, 14 Oct. 

1970. 

29. MeRuer, D., "Human Dynami.cs in Man-Machine Systems," Automatica, 

Vol. 16, May 1980, pp. 237-253. 

30. Jex, H. R. and R. W. Allen, "Research on a New Human Dynamic 

Response Test Battery," 6th Annual Conference on Manual Control, 7-9 Apr. 1970. 

TR-l19;I-l R-3 
This Page Intentionally Left Blank 
APPENDIX A 

DERIVATION OF NIPIP ESTIMATION EQUATION 

FOR THE SHALLOW GLIDE AND FLARE MODEL 

Simplifying notation 

Yp (s) 

e 

6(s) 

(1/Tr)( 1/ ~T) 

Ke(v)e-kTS 

(a)(i3) 

• • -kTs KhKe(v)[hc - h]e 

(O)(a)(i3) 

Ke(v)ee-kTS 

(a)(i3) 

From the Z transform table, Smith, J. A., Mathematical Modeling and 

Digital Simulation for Engineers and Scientists, John Wiley & Sons, NY, 

1977, p. 100. 

6(Z) 

TR-1197-l A-·l 

(h• - h·)Z-k c 
Z-1 

l'1ul tiplying through by -=r and establishing a common denominator 

Z 

6(Z) 

(a_8)V(1_e-aTZ-l)(1_e-8Tz-1) 

+ 8(v-a)(1-z-1)(1-e 8TZ-1) 

+ a(8-v)(1-Z-1)(1-eaTz-1) 

(1_z-1)(1_e-aTz-1)(1_e-STz-1) 

bo + b1Z-1 + b2Z- 2 

-1 -2 -3 l-alZ azZ a3Z -1 -2 -3 l-alZ a2Z - a3Z 

Multiplying through by the denominator leads directly to the estimation 

(difference) equation for hc = 0 

. . . + cohn-k + clhn-1-k + c2hn-2-k + B 

TR-1l97-1 A-2 
APPENDIX B 

SPACE SHUTTLE FLYING QUALITIES QUESTIONNAIRE 

RespoI1ldent's Name -----

Flight ______________ on (date) _________ _ 

The Shuttle Orbiter~ as a large highly augmented, fly-by-wire, delta 

glider has some entry flying characteristics which are considerably different from more conventional aircraft. The current flying qualities 

criteria data base is drawn from experience with the latter and may not 

be appropriate for Shuttle-like vehicles. This questionnaire has been 

prepared to obtain information on manual control of Shuttle entry for 

development of improved criteria for advanced aircraft. 

This questionnaire has two sections, the first concerns the crews' 

perception of the task and flying techniques in each entry flight segmente The second section concerns the crews' evaluation of Orbiter 

dynamics, flying qualities and workload. 

In both sections, five flight segments are of interest: 

1. Supersonic bank reversals (M = 18, 9, 5, 2.5) 

2. Heading Alignment Circle (HAC) turn 

3. Descent on the steep glide slope 

• auto speed brake 

• manual speed brake 

4. Preflare 

5. Shallow glide and final flare to touchdown 

TR-1l97-1 B-1 
SECTION I 

QUESTIONS ON TASKS AND FLYING TECHNIQUE 

As noted on the previous page, the Shuttle Orbiter has some flying 

characteristics which are considerably different from conventional aircraft, not the least of which is the lack of thrust or power to maintain 

constant energy flight segments and/or extract the vehicle from hazardous flight conditions. The questions contained in this section are 

intended to help us establish 

a) the control loop structure in use by the pilot (e.g., what 

parameter is being controlled, what is being tracked) 

b) the pilot control strategy (e.g., constant 6, g, ~, y, ••. ; pitch 

attitude; maintain _?_ within ± ~ switch to another strategy 

when _?_) 

c) pre-programmed (highly learned, open-loop) maneuvers 

We will also use this information in analyzing data recorded during 

your flight (or simulation). For example, it will assist us in 

a) identifying the different flight segments on the flight traces 

b) modeling the pilot's control loop structure and strategy so as to 

extract (via parameter identification techniques) information 

concerning pilot workload, pilot adaptive time to each control 

task or flight segment, etc. 

c) identifying influence of the Orbiter nonconventional flying and 

control characteristics 

d) preparing criteria and design guides for improving the flying 

qualities of future Shuttle-type aircraft. 

Please feel free to add any comments, suggestions, criticism, or 

whatever which you feel may be of assistance in achieving the above 

goals. Your cooperation is greatly appreciated. 

TR-1l97-1 B-2 
A. General questions for each flight segment: 

1 • Please check the control channels flown manually in each flight 

segment 

PITCH ROLL/YAW SPEED BODY 

BRAKES FLAP 

Supersonic 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) 

I 

I 

I 

HAC 

Turn I 

I 

Steep 

Glide 

Preflare 

! 

+---

Shallow 

Glide and 

Final Flare 

TR-1l97-1 B-3 
2. What were the primary objectives of each flight segment flown 

manually, the primary variables controlled (tracked) and the allowable range of these variables? 

Supersonic 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) 

HAC 

Turn 

Steep 

Glide 

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

Preflare 

-

Shallow 

Glide and 

Final Flare 

TR-1197-1 B-4 
3. What cues and conditions are used to transition from each flight segment to the next, what information sources are used, what controller 

inputs are employed? 

-==============~==================~=============-=--=-~.==============-===-========= 

Supersonie 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) I 

HAC 

Turn 

Steep 

Glide 

I 

--.--.. --... -~-.. --.--.-+----------.--------~------~---~---------i 

Preflare 

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

Shallow 

Glide and 

Final Flare 

1'R-1197-1 B-5 
4. To what extent is each flight segment an "open loop" (learned) 

maneuver? To what extent is "closed loop" tracking significant and 

what responGe variables from what sources are involved? 

Supersonic 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) 

HAC 

Turn 

Steep 

Glide 

Preflare 

Shallow 

Glide and 

Final Flare 

TR-1l97-1 

I 

I 

I 

I 

I 

I 

I 

B-6 
5. To what extent are manual trim (pitch, roll, yaw, body flap) controls 

used? 

-

Supersonic I 

I 

Bank i 

Reversals I 

(Please i 

I 

identify I 

I 

specific I maneuvers) I 

I 

, 

i 

! ________ -1- -

i 

HAC I 

Turn I 

I 

I 

I 

I 

-1------

Steep 

Glide 

-------tPreflare i 

---.--~-

Shallow I 

I 

Glide and i 

I 

Final Flare I 

TR-1l97-1 B-7 
6. To what extent did the commander operate in a "head-up" visual mode, 

what factors determined "head-up" operation? 

t---

Supersonic 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) 

----

------------------r--------- --------------------------------------------

I HAC i 

Turn 

---_ ----- ----- ------+-------------------------------1 

Steep 

Glide 

Preflare 

Shallow 

Glide and 

Final Flare 

TR-1197-1 

I 

I ----------

B-8 
7. What is the extent of occupation with secondary axes of control? 

(e"g., heading control in Preflare, etc.). 

SupenlOnic 

Bank 

Reversals 

(Please I 

I ident:lfy 

specific I 

maneuvers) ! 

! 

i 

I 

- I 

I 

I 

I 

HAC I 

Turn 

I 

, 

i 

! 

! 

i 

I 

I 

Steep f 

! 

Glide 

I 

Preflare I 

I 

Shallow 

Glide and 

Final Flare 

I 

TR-1l97-1 B-9 
8. What (if any) disturbances (crosswinds, turbulence) were significant, 

what techniques were used to regulate against these disturbances? 

Supersonic 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) 

HAC 

Turn 

Steep 

Glide 

Preflare 

Shallow 

Glide and 

Final Flare 

TR-1l97-1 B-IO 
B. Specific Questions for Individual Manually Controlled Flight Segments 

1. Steep glide slope acquisition and glide 

a. To what extent is manual control of speed with speedbrakes 

similar to conventional use of throttles? 

b. Is speed brake operation basically "open loop" or "closed 

loop," continuous or discrete? 

c. Are there any significant interactions between manual (or 

automatic) speedbr~ke control and attitude/path control with 

the RHC? If so, how does this effect piloting technique? 

2. Preflare Pullups 

a. What cues are used to initiate preflare? 

b. Do you prefer to fly this 

as a constant load factor maneuver? Why? 

as a constant pitch rate maneuver? Why? 

other? Why? 

TR-1l97-1 B-11 
c. Any comments as to the RHe force/displacement/sensitivity/ 

lat-long harmony during this maneuver? 

d. What determines speedbrake retraction? 

e. What determines gear extension, is the gear light used? 

3. Acquisition of and glide on the shallow glides lope and final 

flare 

a. Is the shallow glide basically constant sinkrate, constant 

flight path angle or neither, how is the controlled variable 

maintained? 

b. Is there 

open or 

flare? 

a distinct final flare maneuver? Is it basically 

closed loop? What determines the initiation of 

Is flare due to ground effect noticeable or 

explicitly accounted for? 

c. Is there any conscious effort to "schedule" sinkrate with 

altitude in this flight segment? 

d. What are the relative priorities and criteria in the control 

of: energy, touchdown sinkrate, touchdown speed, touchdown 

point? 

TR-1197-1 B-12 
e. What are the principal attitude and path references? What is 

the relative use of "out the window," HUD and panel displayed 

information? How are verbal callouts from other crew members 

used? 

f. Is the selection of the shallow glides lope and flare characteristics essentially preplanned or are they modified as the 

landing unfolds? Would the strategy be modified in 

crosswinds? 

g. What situations might lead to ballooning? To a pilot induced 

oscillation (PIO)? 

h. What procedures are used to control (touchdown) speed? 

i. Is attitude control basically open or closed loop? Are attitude commands basically discrete or continuous? 

j. Are you conscious of any pulse type RHC control activity? 

Does this seem "natural"? necessary? 

TR-1197-1 B-13 
SECTION II 

A. General Evaluation of Orbiter Dynamics, Flying Qualities and Workload 

The following is a list of flying quality related characteristics 

which mayor may not have adversely impacted manual control workload, 

task difficulty, attitude or path control precision, etc. These have 

not been integrated into a question format in order to avoid 

restricting the nature of your response. Comments are therefore 

solicited on any aspect in which a particular factor may stand out in 

your memory as adversely impacting the above during the flight segments of interest. Please identify the flight segment being commented on. 

1. Longitudinal control 

a. Rotational hand controller (RHC) characteristics 

RHC displacements 

RHC force gradient 

Breakout sensitivity 

Gain (rate command/force input) and shaping 

Longitudinal-lateral displacement, gradient, or sensitivity harmony 

b. Pitch attitude response (to inputs required to perform task) 

Effective time delay 

TR-1197-1 B-14 
Initial response onset (rise time) 

Overshoot 

Settling time 

Predictability 

Sensitivity 

PIO tendency 

c. Path response/control 

Effective motion delay time 

Predictability 

PIO tendency 

Any special control techniques employed? required? 

d. Airspeed control 

Precision 

Predictability 

TR-1197-1 B-lS 
e. Disturbances 

Turbulence 

Wind shear 

Ground effect 

f. Workload 

Is control workload significant? dominant? 

Does other task workload detract from control task 

performance? 

Cooper-Harper rating (if possible) 

2. Lateral-Directional Control 

a. Rotational hand controller characteristics 

RHC displacements 

RHC force gradient and shaping 

Breakout sensitivity 

Lateral-longitudinal harmony 

TR-1l97-1 B-16 
Gain (rate command/force input) and shaping 

b. Roll attitude response 

Effective time delay 

Initial response onset (rise time) 

Overshoot 

Settling time 

Predictability 

Sensitivity 

PIO tendency 

Lateral acceleration at pilot 

Roll ratcheting 

c. Heading response/precision 

Roll into turns 

TR-l197-1 8-17 
Rollout of turns 

d. Workload 

Is control workload significant? 

Does other task workload detract from control task? 

e. Cooper-Harper rating (if possible) 

3. Summary (Brief) 

a. Major problems 

b. Good features 

TR-1197-1 B-18 
8. Specific Questions on Flying Qualities and Workload 

1. In landing, are there any unusual characteristics of the Shuttle 

pitch attitude response to the rotational hand controller? Does 

the response appear to be "rate command," "attitude command" or 

neither? 

2. Are there any unusual characteristics of the Shuttle path 

(altitude, flight path angle) response to pitch attitude changes? 

3. What differences in pitch trim and airspeed control, as compared 

to conventional aircraft, are requi~ed because of the zero stick 

force/speed gradient of the Shuttle's pitch rate command system? 

4. What is the relative difficulty of speed control with the speedbrakes (steep glideslope)? 

5. In the shallow glide and final flare are there any conflicts in 

simultaneous control of touchdown point, speed and sinkrate? 

TR-1197-1 B-19 
6. Are there inadequacies in the availability of information (pitch 

attitude, sinkrate, etc.)? What display changes would help? How 

adequate is the view "out the window" for attitude, altitude, 

heading, and flight path control? 

Supersonic 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) 

HAC 

Turn 

Steep 

Glide 

Preflare I 

Shallow 

Glide and 

Final Flare 

TR-1l97-1 

I 

I 

I 

B-20 
7" Do you foresee any operational conditions (turbulence, crosswinds, nigllt landing, etc.) which might approach flying qualities 

limits? What response characteristics of the Shuttle might be 

limiting in these situations? 

8. To what extent did the actual Shuttle Orbiter flying characteristcs in approach and landing differ from ground simulations and 

STA flights? What, if any, changes would be valuable for preflight ground and STA simulations? 

TR-1197-1 B-21 
9. What portion (rough percentage) of the total crew workload capacity was used in each flight segment? 

Supersonic 

Bank 

Reversals 

(Please 

identify 

specific 

maneuvers) 

-

HAC 

Turn 

-----

Steep 

Glide 

Preflare 

-tShallow I 

Glide and 

Final Flare 

TR-1l97-1 

NO 

WORKLOAD 

0 

I --I 

I 

I 

I 

I 

I 

r 

i 

I 

I 

i 

I 

I 

B-22 

LIMIT OF 

CAPACITY 

50% 100% 

I 

I 

I 

i 

I 
I - 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. 

NASA CR-170407 

4. Title and Subtitlll 5. Report Date 

Sp,ace Shuttle Flying Qualities Criteria Assessment - Phase III February 1984 

6. Performing Organization Code 

7. Author(s) 8. Performing Organization Report No. 

T. T. Myers,. D. E. Johnston, and D. T. McRuer TR-1197-1 

10. Work Unit No. 

9. Performing Organization Name and Address 

Systems Technology, Inc. 

13766 South Hawthorne Boulevard 11. Contract or Grant No. 

NAS4-3005 Ha'~thorne , California 90250 

13. Type of Report and Period Covered 

12. Sponsoring Agency Name and Address Contractor Report - Final 

National Aeronautics and Space Administrat:ion 

Washington, D.C. 20546 14. Sponsoring Agency Code 

RTOP 506-63-40 

15. Supplementary Notes 

NASA Technic:al Monitor: Donald T. Berry, Ames Research Center, Dryden Flight Research Facility, 

Edl~ards, CA 93523. Phase I of this study is published as NASA CR-170391 ; phase II, as 

NASA CR-170~106. 

16. Abstract 

-

This study continues the development of flying qualities experiments as part of the Orbiter Experiments Program (OEX). Phase I, 

published as CR-170391, reviewed flying qualities criteria and initial Shuttle data. Phase II, published as CR-170406, proposed a 

program implementation conSisting of inflight experiments correlated 

with extensive simulator programs, a preliminary review of instrumentation and data analysis requirements and a review of STS 1 

through 4 flying qualities. This report explores the crucial flight 

data measurement and reduction techniques required for the experimental approach. An overvi.ew of available flight data is presented, 

identification of the effective augmented vehicle and pilot models 

determined, a summary of flight data problems compiled, and further 

recommendations for the Orbiter Flying Qualities (OFQ) experiment 

provided. 

17. Key Words (Suggested by Author(s)) 18. Distribution Statement 

Space shuttle orbiter Unclassified-Unlimited 

Flying qualities 

Fli.ght control 

Pilot models STAR category 08 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price* 

Unclassified Unclassified 128 A07 

*For sale by the National Technical Ir~ormation Service, Springfield, Virginia 22161. 
End of Document