NASA/CP-2001-210389/VOL3

Pilot-Induced Oscillation Research:

Status at the End of the Century

Compiled by Mary F. Shafer and Patti Steinmetz

NASA Dryden Flight Research Center

Edwards, California

April 2001
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NASA/CP-2001 =210389/VOL3

Pilot-Induced Oscillation Research:

Status at the End of the Century

Compiled by Mary F. Shafer and Paul Steinmetz

NASA Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273

April 2001
NOTICE

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Space Administration.

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Foreword

"Pilot-Induced Oscillation Research: The Status at the End of the Century," a workshop

held at NASA Dryden Flight Research Center on 6-8 April 1999, may well be the last

large international workshop of the twentieth century on pilot-induced oscillation (PIO).

With nearly a hundred attendees from ten countries and thirty presentations (plus two that

were not presented but are included in the proceedings) the workshop did indeed

represent the status of PIO at the end of the century.

These presentations address the most current information available, addressing regulatory

issues, flight test, safety, modeling, prediction, simulation, mitigation or prevention, and

areas that require further research. All presentations were approved for publication as

unclassified documents with no limits on their distribution.

This proceedings include the viewgraphs (some with authors' notes) used for the thirty

presentations that were actually given as well as two presentations that were not given

because of time limitations. Four technical papers on this subject that offer this

information in a more complete form are also included. In addition, copies of the related

announcements and the program are incorporated, to better place the workshop in the

context in which it was presented.

Mary F. Shafer

.°.

111
CONTENTS

°oo

FOREWORD ..................................................................... 111

VOLUME 1 - SESSIONS I - III

SESSION I - 6 APRIL 1999 .......................................................... 1

1. Modeling the Human Pilot in Single-Axis Linear and Nonlinear Tracking Tasks

Yasser Zeyada and Ronald A. Hess, University of California, Davis ........................... 3

2. Bandwidth Criteria for Category I and II PIOs

David G. Mitchell, Hoh Aeronautics, Inc; and David H. Klyde, Systems Technology, Inc .......... 17

3. Criteria for Category I PIOs of Transports Based on Equivalent Systems and Bandwidth

Kenneth F. Rossitto and Edmund J. Field, Boeing Phantom Works ........................... 29

4. Designing to Prevent PIO

John C. Gibson, Consultant, British Aerospace ........................................... 41

SESSION II - 6 APRIL 1999 ........................................................ 53

5. Replicating HAVE PIO on the NASA Ames VMS

Jeffery Schroeder, NASA Ames Research Center ......................................... 55

6. Replicating HAVE PIO on Air Force Simulators

Ba T. Nguyen, Air Force Research Laboratory ........................................... 65

7. Prediction of Longitudinal Pilot-Induced Oscillations Using a Low Order

Equivalent System Approach

John Hodgkinson and Paul T. Glessner, Boeing; and David G. Mitchell, Hoh Aeronautics, Inc ...... 67

8. Recommendations to Improve Future PIO Simulations

Brian K. Stadler, Air Force Research Laboratory .......................................... 81

SESSION III - 7 APRIL 1999 ....................................................... 93

9. FAA's History with APC

Guy C. Thiel, FAA ................................................................. 95

10. PIO and the CAA

Graham Weightman, JAA (UK CAA) ................................................. 107

11. PIO Flight Test Experience at Boeing (Puget Sound) - and the Need for More Research

Brian P. Lee, Boeing Commercial Airplane Group ....................................... 109

12. The Effects on Flying Qualities and PIO of Non-Linearities in Control Systems

Edmund Field, Boeing Phantom Works ................................................ 157

13. Mitigating the APC Threat - a work in progress

Ralph A'Harrah, NASA Headquarters ................................................. 171

V
VOLUME 2 - SESSIONS IV - V

SESSION IV - 7 APRIL 1999 ...................................................... 179

14. Flight Testing for APC: Current Practice at Airbus

Pierre Poncelet, Aerospatiale Aeronautique; and Fernando Alonso, Airbus Industrie ............. 181

15. The Prediction and Suppression of PIO Susceptibility of Large Transport Aircraft

Rogier van der Weerd, Delft University of Technology .................................... 189

16. Flight Testing For PIO

Ralph H. Smith, High Plains Engineering .............................................. 217

17. Use of In-Flight Simulators for PIO Susceptibility Testing and for Flight Test Training

Michael Parrag, Veridian Engineering ................................................. 227

18. A Method for the Flight Test Evaluation of PIO Susceptibility

Thomas R. Twisdale and Michael K. Nelson, USAF Test Pilot School ........................ 259

SESSION V - 8 APRIL 1999 ....................................................... 269

19. Onboard PIO Detection and Prevention

David B. Leggett, Air Force Research Laboratory ........................................ 271

:20. Real Time PIO Detection and Compensation

Chadwick J. Cox, Carl Lewis, Robert Pap, and Brian Hall, Accurate Automation Corporation ..... 279

21. PIO Detection with a Real-time Oscillation Verifier (ROVER)

David G. Mitchell, Hoh Aeronautics, Inc ............................................... 299

22. Pilot Opinion Ratings and PIO

Michael K. Nelson and Thomas R. Twisdale, USAF Test Pilot School ........................ 305

23. The Need for PIO Demonstration Maneuvers

Vineet Sahasrabudhe and David H. Klyde, Systems Technology, Inc.;

and David G. Mitchell, Hoh Aeronautics, Inc ............................................ 307

VOLUME 3 - SESSION VI AND APPENDICES

SESSION VI- 8 APRIL 1999 ...................................................... 317

24. Boeing T-45 Ground Handling Characteristics

James G. Reinsberg, Boeing St. Louis ................................................. 319

25. Extraction of Pilot-Vehicle Characteristics from Flight Data in the Presence

of Rate Limiting

David H. Klyde, Systems Technology, Inc.; and David G. Mitchell, Hoh Aeronautics, Inc ........ 329

26. Comparison of PIO Severity from Flight and Simulation

Thomas J. Cord, Air Force Research Laboratory ......................................... 343

27. A Summary of the Ground Simulation Comparison Study (GSCS)

for Transport Aircraft

Terry yon Klein, Boeing Phantom Works .............................................. 359

28. Real Experiences in the Frequency Domain

Randall E. Bailey and Andrew Markofski, Veridian Engineering ............................ 365

29. Pilot Modeling for Resolving Opinion Rating Discrepancies

David B. Doman, Air Force Research Laboratory ........................................ 391

vi
30. Closing Remarks

Mary F. Shafer, NASA Dryden Flight Research Center ................................... 397

APPENDIX 1: ANNOUNCEMENTS, INFORMATION, AND PROGRAM ................ 399

1. Announcement and call for papers ............................................... 401

2. Information for presenters and attendees (sent by e-mail) ............................. 403

3. Program ................................................................... 409

APPENDIX 2: PRESENTATIONS PRINTED BUT NOT GIVEN AT THE WORKSHOP... 411

1. Recent Results of APC Testing with ATTAS

Holger Duda and Gunnar Duus, Deutches Zentrum fiir Luft- und Raumfahrt e.V ............... 413

2. Criteria to Simulation to Flight Test--and Vice Versa

David G. Mitchell, Hoh Aeronautics, Inc ............................................... 439

APPENDIX 3: PAPERS SUPPORTING WORKSHOP PRESENTATIONS ............... 453

1. Designing to Prevent Safety-Related PIO

J.C. Gibson, British Aerospace Warton (retired), Consultant ................................ 455

2. Pilot-lnduced Oscillation Prediction with Three Levels

of Simulation Motion Displacement

Jeffery A. Schroeder, William W.Y. Chung, and Duc T. Tran, NASA Ames Research Center;

and Soren Laforce and Norman J. Bengford, SYRE Logicon ................................ 475

3. A Method for the Flight Test Evaluation of PIO Susceptibility

Thomas R. Twisdale and Michael K. Nelson, USAF Test Pilot School ........................ 487

4. Pilot Opinion Ratings and PIO

Michael K. Nelson and Thomas R. Twisdale, USAF Test Pilot School ........................ 493

vii
Session VI

317
q
T45TS Boeing T45 ;J

Ground Handling Characteristics ---_ _.

NASA Dryden Workshop

Jim Reinsberg

Principal Technical Specialist

T45TS Aerodynamics, Flying Qualities

The Boeing Company

(314)233-1092

james.g.reinsberg@boeing.com

6-8 Apr 99

 T4 sTS T45 Aircraft Description

Derived from BaE Hawk

Typical Weight Data: //"-/

_ Eax t_e, load, 2 ..... 13,381 / / /

39.33 fl _1

Key aircraft components:

> -12% of weight on nose landing gear

> Single chambered, semi-levered main landing gear

> Single chambered, cantilevered nose landing gear (2 tires)

> 20 deg/sec nose wheel steering (NWS) - 12 deg deft max

> Reversible, mechanical rudder

> Hydraulic powered aileron, stabilator.

> Limited Yaw Damper Control (YDC)

6-8 Apr 99

319
T45TS Summary of T45 "J

Ground Handling Issue __ -

Directional control issues have been with the T45 since

1989. This is a basic airframe issue. Multiple "Triggers" such as

cross-winds, inadvertent brake/NWS/rudder inputs, blown tire,

aggressive corrections, etc. create a control problem which is

amplified by "Sustainers" such as landing gear dynamics, brake

sensitivity and feel, roll/yaw coupling, lateral acceleration cues, etc.

Over the years many attempts and studies have been undertaken to

improve basic airframe handling characteristics with some success.

But fixes are not easy or "cheap". The lack of a good ground

handling METRIC has dampened the enthusiasm to flight test

"potential fixes".

6-8 Apr 99

BAJUSN

Efforts Toward Resolution

Solutions Investigated With Mixed Success

Nov 89

Nov 90

May 93

Dec 93

Mar 94

Mar 94

Jun 94

Sep 94

Nov 95

Jan 97

Aug 98

Established SA-4A during DT-IIA:

- =Directionalpilot induced oscillations during landingrollout."

Developed current production NWS system

- Full time NWS cleared "PIO"yellow sheet SA-4A

- EnteredFleet Aug 92

Established SA-162 during DT-II:

- =Overlysensitivedirectional controlcharacteristicsduring landingrollout,"

Developed 1st industry ground handling PIO metric

- Provideda =yardstick"for predictingeffectivenessof modifications

ADR data @ KNAS supported PIO metric

Started flight evaluation of higher rate NWS system

- ImprovedhandlingbutPIO susceptibilityremained

Joint USN/MDA "PIO team" formed to explore causes and solutions

Recommended fix of high gain yaw damping with higher rate NWS

Started flight evaluation of "PIO team" recommended fix

- Concludedimprovements not adequatefor production

- Identifiedobjectionablegroundhandlingotherthan PIO

NAVAIR recommended assessment by outside company

Started independent assessment with STI, subvendor to BA

6-8 Apt 99

320
Boeing Criteria for ,.J

T45TS Ground PIO Susceptibility "-'__

• Applied Mil STD criteria for longitudinal PIO (Ralph Smith).

- Showed this to be a good predictor of directional PIO tendencies with:

> Frequency response of flight test data

> Six degree of freedom (6-DOF) analysis with 0.25 sec time delay pilot modet

• MDA experience at this time:

- 10 PA landings were analyzed - included a variety of pilots, crosswinds, and braking tasks.

• Ny at pilot and yaw rate (R) considered most significant control parameters

• Bode plots: 0.6 Hz control from Ny feedback, 1.0 Hz control from R feedback

- A015 landing rollout PIO shows pilot "responding" to Ny

• Criteria successfully predicted higher rate NWS would not reduce PIO potential.

• Employed as metric for joint USN/Boeing PIO Susceptibility team

- Goal: Achieve F-18 Ny phase response.

- Identified 50 potential causes. 8 most promising showed no single or combined root cause.

- Analyzed 3 augmented control solutions:

• R + Ny feedback to NWS, R command, and R feedback to rudder

• R feedback to rudder met F-18 Ny phase criteria.

I Improved, high rate PWM NWS and YDC-10 approved for flight test. I _ _,99

Results Of YDC-10

Flight Test Program

Steering Control Electronic Set (SCES) 1.4

• Allowed testing of production and "test" software with a bit flag change.

• Production T45 NWS software:

- Bang-bang controller, 20 deg/sec max no-loed rate

- Turn-on at 0.75 deg error, rum-off at 0.5 deg error.

- Low gain steering: linear slope, 2.5 inches of pedal -> 12 deg of NWS

• Pulse Width Modulation (PWM) software:

- Still a bang-bang controller, but

• 5 discrete no-load rates, from 8 deg/sec to 52 deg/sec

• Uses •look-ahead" to determine best control speed

• Narrows turn-on/turn-off threshold when pedals moving

• Variety of pedal -> NWS schedules available

NOTE: PWM also required a hydraulic supply orifice change to achieve higher

no-load rate.

6-8 Apr _

321
T45TS Results Of YDC-10

Flight Test Program

Centedine Crossing Task

,_ I 1000'172_ if500ft I 1000ft I , 5ffAdeqJate

Oossing/_e _ 2000 f_

CROSS

- Low gain and low predictability

- Significant variations in crossing angle

- YDC tends to washout initial input

RE-ACQUI SITION

- High gain, high accelerations/rates

- Susceptible to "roll/yaw"

- Steeper x-ing angle, harder task, prone to centedine overshoot

TRACK

- High gain, low Ny, moderate yaw rate

- Performance degraded if Phase 2 overshoots desired criteria

Combined with other variations

(weight, crosswind, inadvertent

differential braking), significant

run-to-run variations in task

difficulty can occur.

6.8 Apt g9

°T45TS Results Of YDC-10 -J

- - Flight Test Program -_:_-_-

• FREQUENCY DOMAIN ANALYSIS

- Predicted reductions in Ny phase lag were achieved

> Only for small inputs (-25%) due to yaw damper saturation

- High rate NWS had no effect on Ny or R phase lag

- Centedine x-ing maneuver did produce PIOs during Re-acquisition and Tracking

> ONLY with non-optimum YDC feedback gain

> Re-acquisition PIOs" High Ny -> roll/yaw

> Tracking PIOs: Low Ny -> often ignored in pilot comments

• PILOT COMMENTS

- PIO ratings slightly reduced with YDCJPWM.

- Significant factors other that phase lag influencing the pilot:

> Velocity vector loosely coupled to nose

> Roll opposite yaw - =leans"

> Inadvertent NWS inputs

> Insufficient brake pedal (force) feedback

> Rudder pedal mechanical characteristics

> Crosswinds

CONCLUSIONS: Incremental improvement for small pedal inputs only, and would

not close yellow sheet SA-162. r_ _o,

322
T45TS Results Of YDC-IO

Flight Test Program

ROLL _NG - _G

p ,,-

i :

_ -4---- 4--

.... i-L ---i--

-_ ::--

_ _ "

-T" ..--r-

....i--_"i....

"i .... _-

NY PILOT - G'S YAW RATE - D/S RUD PEDAL - fN NWS DEFt. - DEC

,

--.4-- _ -......-_ _.4----4

.--'---H

-T-

"!_II

,,_

,.._.. .._...._

f..-i.w,:

_---

p__:_.__

• --_ ._ .....

7'

D_

|

"4"i_

3_Y

_z z

6-8 Apt 99

Results Of YDC-IO

Flight Test Program

ROLL*I_WG- DEG NY PILOt" - G'S YAW RATE - D/$ RUD PEDAL - IN NWS DEP"L- DEG

i _ --.

Q_

M _z>

6-8 h,pr 99

323
T45TS NASA LaRC Analysis

of T45 Tires

• METHOD:

- Used Low speed Tire Test Vehicle (LTTV) to measure cornering performance of nose and main

tires under full scale, realistic surface conditions.

> Max vertical load 6000 Ib

> Max tire yaw angle 90 deg

> Max speed 60 mph

- Varied tire pressure (field, carrier), vertical load and skid angle.

- Nose tire is very under-loaded at 300-900 Ib per tire (5-6% vs. design 32%).

- LTTV data validated by flight test trajectory matching.

• CONCLUSION:

- Main tire cornering stiffness less than modeled by 13-44°/o, depending on normal load.

- Main tire cornering stiffness reduction with normal load more than currently modeled.

- Nose tire cornering stiffness more than modeled by 6-19%, depending on normal load.

I A ground handling assessment REQUIRES accurate tire data under realistic surface _l

conditions. The LTTV proved to be a rapid and economical tool for gathering T45 tire data.

Other NASA facilities exist for tires with greater vertical Ioadings. I

6-8 Apr 99

T45TS

Independent Assessment

Contract With STI

• Objective and Product:

- Analytical assessment by Systems Technology Incorporated (STI)

- Recommend procedures and/or aircraft modifications with the potential to

minimize or eliminate undesirable landing rollout characteristics.

- Feasible recommendations will likely require additional research and flight

evaluation by USN/BA team prior to production consideration

• Tasks:

• Status:

- Review past efforts

- Examine basic aircraft design issues

- Recommend a way forward

7 Feb 98 USN issued RFP to Boeing (BA)

21 Apr 98 BA selected STI as winning subvendor

21 Jul 98 USN/BA complete contract negotiations

20 Aug 98 Kickoff meeting in STL. BA, STI & NAVAIR (15 month contract)

16 Nov 98 First quarterly review

18 Feb 99 Second quarterly review

15-19 Feb 99 - First flight simulation

S-SA_ 99

324
T45TS

Independent Assessment

Contract With STI

Status After First Flight Simulation

• NASA LARCtire data incorporatedintoall 6-DOF models.

• Analysis offlight test data suggestthat heading,an.qlefeedback is the primary pilot

control mechanism.

• Boeing6-DOF and STI linearmodel have been benchmarked to flight test data.

• STI Linear model analysis shows that the T45 -

- has an oversteer characteristic (tire cornering stiffness is key)

- has a critical speed, above which the vehicle has an unstable pole (- 60 kts).

The understeergradientUG may be a reliablemetricfor PIO potential

UG = 32.17*57.3*{(m/t)*[(b/Y_f) - (a/Y_r)]} [deg/g]

m = vehicle mass [slugs]

a = distance from front tire to cg [ft]

b = distance from rear tire to cg [ft]

I = distance from front to rear tire (l=a+b) {ft]

Y(,f = front axle "aero+tire+.." cornering coefficient [llY/rad]

Y_, = rear axle "aero+tire+.." cornering coefficient [Ibf/rad] 6-8 Apt 99

T45Ts

Independent Assessment

Contract With STI

Status After First Flight Simulation

Maneuvers used during first simulation:

- Constant radius rum circle (2000 ft)

- Maximum heading capture and stabilization (aggressive)

- Heading capture and hold (instruments only - no visual)

- Heading angle sum-of-sines tracking (instruments only - no visual)

- Runway centerline tracking with crosswind gust disturbance

• Aircraft parameters varied during first simulation:

- Fuel (empty, 65% full)

- Aircraft understeer gradient, UG

- Nose wheel steering actuator model (production and =ideal")

Preliminary findings:

- Fixed base simulation: not perfect, but we're working on it

- "Ideal" actuator model: most effect on fine tracking, not PIO

- Turn circles show a break in roll vs. Ny at 0.2 g's (approx 2 deg roll)

- HQR and PIO ratings track understeer gradient UG

- A 2 point HQR/PIO reduction may be possible with a tire change

6-8 Apr g9

325
Independent Assessment ..

T45TS Contract With STI "-,_'_,*

Status After First Flight Simulation

Excellent agreement between flight test, flight simulation and Boeing 6-dof (MODSDF)

O. ! _IdlSec| iO tOO

4o 1 .......................... _ . i ........................ I

_ . :...i : i } Yaw Rate to Pedal ] "

Gain, dB _ _._= -:-.................................. :

&

Phase, deg o'- -; :-_ .... ?.:: ....:

.tO0=_

-2OO oT-_ Fillet fit W Iltl Vl IllU)ltl_ MIII_I Ill fit to = tire tilt!

M|I / t TIl_k 4 Sll

Low Power i

in

Flight Test Data

6-8 A_r 99

Independent Assessment 1

:.T.45-/T:S Contract With STI _4_:_*:___. :* ....

Status After First Flight Simulation

From flight test: More than 2 deg of roll was consistently remarked as "very uncomfortable".

Below 2 deg of roll, it was otlen ignored.

5

, _.a_/u [From flight simulation [

, _ _ I,,,,-r, circletosts__________L_l

(¢Jegl

O_

-0.5 -0.4 -0.3 -0.2 -0.1 0

N),_ (g's) 6-8 Apr 99

326
T45TS

Independent Assessment

Contract With STI

Status After First Flight Simulation

l:I::::= r--r-'!-:r-.!

....+_,....;---i-_-_I

.!.i_!!ii.!,!_,!

Heading Capture and Hold: t :t_-'_'._ !

> I0 deg heading change

:,k-,.-7-i i_ri_ /

...... .,.a.' ',.,.',a, ° ' °

Runway Ccn:e.rhne Tracking 't --'T;-" F-i -T T I-'l

> randomx-winds during tracking _.a.._.-,=M_

p J _ c

6-8 Apr 99

Independent Assessment

Contract With STI

Future Efforts

• Refine Boeing flight simulation

- Adjust seat/pedal/heel-rest to T45 spec

• Pilot-vehicle analysis:

- Acquire flight test data from dissimilar aircraft

- Complete pilot-vehicle analysis of ground handling dynamics:

> Ergonomics (braking, steering crossover)

> Control sensitivity and magnitude

• Crosswinds

Refine tasks/metricsto quantifyexpected improvements

- Define new, or modify existing tasks.

- Quantify possible "improvements" in flight simulation

Present final report/recommendations: November, 99

6-8 Apr99

327
q
EXTRACTION OF PILOT-VEHICLE

CHARACTERISTICS FROM FLIGHT DATA

IN THE PRESENCE OF RATE LIMITING

David H. Klyde

dldyde@sy stemstech.com

Systems Technology, Inc.

David G. Mitchell

Hoh Aeronautics, Inc.

Pilot-lnduced Oscillation Research:

The Status at the End of the Century

NASA Dryden Flight Research Center

6-8 April 1999

PRESENTATION OUTLINE

• Program Overview

• Background

- Category II PIOs

- Airplane Bandwidth/Phase Delay Criteria

• F-14 Dual Hydraulic Failure Flight Test Program

Flight Test Data Description

Flight Test Data Analyses

• Conclusions

8 April 99 PIO Research Status Workshop O

329
PROGRAM OVERVIEW

Illl I I_:.

• Work performed by Systems Technology, Inc.

(STI) under a subcontract from Hoh Aeronautics,

Inc. (HAI)

• Part ofa HAI Phase II SBIR with the Air Vehicles

Directorate of the Air Force Research Laboratory

• Air Force Project Engineer - Thomas J. Cord

• F-14 flight data provided by Naval Air Warfare

Center, Aircraft Division

8 April 99 PIO Research Status Workshop

TIME HISTORY OF THE X-15

LANDING/FLARE PIO

_, fps

v,.,,,,.,,,_[ I I I I I I I_

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Rcf. NASA TN D-1057

8 April 99 PIO Research Status Workshop

330
CATEGORY II PlOs

• Essentially nonlinear pilot-vehicle system

oscillations with amplitudes well into the range

where rate and/or position limits become dominant

• Transitional category between Category I and the

most general, nonlinear Category III PIOs

• Most common jump-resonant, limit-cycle, PIO

event

Intrinsically severe PIO s

8 April 99 PIO Research Status Workshop

8 April 99

CATEGORY II ISSUES

• Presence of rate limiting and other nonlinearities result in a

Frequency and Amplitude dependence

• There are, therefore, a task dependent family of solutions

that will determine PIO susceptibility

• Rate and/or position limiting within a closed-loop structure

will disrupt the aircraft augmentation as the limiter

becomes active

• Criteria will be inherently more complicated in their

application

• Ready applicability of criteria may imply a need for

specific software applications

PlO Research Status Workshop 0

331
8 April 99

CATEGORY II FLIGHT DATA

IIIII I I i ,11,

• All candidate criteria are tentative until val idated with flight

data (qualitative & quantitative)

• Until recently available flight data has been extremely

limited and incomplete (essentially time histories from

flight test of developmental aircraft)

• HAVE LIMITS (USAF TPS Class 96B)

- Configurations flown with variable stability NT-33A

- Reference AFFTC-TR-97-12 (approved for public release)

• USAF TIFS Study

- Parallel HAVE LIMITS with large aircraft configurations

PlO Research Status V_rkshop

BANDWIDTH/PHASE DELAY

REQUIREMENTS

8 April 99 PIO Research Status Workshop

332
BANDWIDTH/PHASE DELAY

,_ I III IIII I$' "rf'r_ _'_ •

• Use flight derived frequency response (nonlinearities

included) to compute Bandwidth (C%w) and Phase Delay

(Xp) parameters for a variety of input amplitude levels

• Assume linear requirements apply to nonlinear (quasilinear) configurations at each input amplitude

• A Bandwidth/Phase Delay locus that is a function of input

amplitude is overlaid on the linear requirements to define

PIO-prone regions

• The input amplitude conditions (Ai) corresponding to the

boundary crossing of the [Xp, 0)Bw](Ai) locus indicates a

critical region for possible onset of Category II PIO

8 April 99 PIO Research Status Workshop

BANDWIDTH/PHASE DELAY

(concluded)

• The transition from a phase margin bandwidth condition to a

gain margin bandwidth condition can be indicative of a Category

II jump resonance phenomenon

• A systematic approach to specify pilot input magnitude for

conducting frequency sweeps is needed

• Drops in coherence occur whenever power is present in the

output that does not correspond to the PVS input, such as pilotinduced noise (remnant), sampling harmonics, and nonlinearities

• Analysis of available data otten indicates a reduction in

describing function coherence in the neighborhood of the onset

or saturation frequency of the rate limiter

8 April 99 PIO Research Status Workshop O

333
DESCRIBING FUNCTION VARIATIONS

WITH INPUT AMPLITUDE

_,_: _,.,, I Ill Ill II I I Ill I II _ _',':,-:_,_"

8 April 99 PIO Research Status Workshop

BANDWIDTH/PHASE DELAY

INPUT AMPLITUDE SENSITIVITY

0.4

0.35

W'_ 0.3

0,25

__ 0.2

a QAS ,,

!

.1: 0.1

G.

O.OS

0

O

! I _ i J

I I I Phase Margin BW |

_----_-[ • Q.,,..rg,..w r---

!1_

! J _L u

q _A-6 o Olin - _.S _ i

1 2 3 4 5

Pitch Attitude Bandwidth, (o 8w_ (radlsec)

8Apd199 PIO Research Status Workshop @

334
F-14 DUAL HYDRAULIC FAILURE

FLIGHT TEST PROGRAM

8 April 99

• Navy flight test program was conducted from 1.0/90 to

3/91.

• The back-up flight control module (BUFCM) was

evaluated for in-flight refueling and landing.

• Maximum stabilator rates were 10 and 5 deg/sec for

BUFCM-HIGH and BUFCM-LOW modes, respectively.

• Aircraft demonstrated good handling in formation flight.

• A number of PIOs were encountered during in-flight

refueling, drogue tracking, and offset field landings.

An excellent PIO database was inadvertently created.

PIO Research Status Workshop

FLIGHT TEST DATA ANALYSES

• Flight Test Data Description

• Example Time Histories

• Identification of Stick Dynamics

• Effects of Rate Limiting

• Identification of PIO Frequency and Task

Bandwidth

• Airplane Bandwidth/Phase Delay Assessments

8 April 99 PIO Research Status Workshop 0

335
FLIGHT TEST DATA

DESCRIPTION

• High quality time history data for:

- 7 frequency sweeps

- 8 drogue hook-ups

- 2 drogue tracking runs

- l field offset landing

• Runs were characterized by:

- Aircraft configuration: wing sweep, gear and flap positions

- Flight condition: altitude, airspeed, Mach number

- FC mode: SAS On, SAS Off, BUFCM-HIGH, BUFCM-LOW

8 April 99 PIO Research Status V_rkshop

BUFCM-HIGH FREQUENCY

SWEEP TIME HISTORIES

40. tl.

i  ̧ iiii_ii_i_:'=_ll_"

_.| *"lll_lIWIIf[

-10t

° 10 20 30 40 N N 70 H N

Time (=eel

8 April 99 PIO Research Status Workshop

336
BUFCM-HIGH DROGUE

TRACKING TIME HISTORIES

4O1

" , I h, ,

=-20[L I'

1°.

,°t

,4 :rl:,

-q0 '

-16 0 10 ' ' 20 30 40 _ SO

8 April 99 PIO Research Status Workshop

BUFCM-HIGH DROGUE HOOK-UP

TIME HISTORIES

40

30

_0

_°

._ .1o

=.20

.,lo

r

1rime loot)

lO

o

==

-10

-1°

,. ,_ .

i'

i lO _l 3o 20 20

(°ec)

8 April 99 PIO Research Status Workshop

337
0

"E

.=.

-40

LONGITUDINAL

STICK DYNAMICS

400.1 t t0 "2000.1

= FLO,v = d+5.46H+l13.15

1 1o

t I

• o+| •

<-I 0.2 _

t 10

Frequency (mcl_ec)

8 April 99 PIO Research StatusWorkshop

EFFECTS OF RATE LIMITING

ON q/FLo N

lo

I= • •

3_ -;to •

-4O

1

8 April99

• i_ oa, uoo*.o_

• I_ o¢I,1,4¢oli._,ii

• IIINcll Xlgl_ i,i_

• l_Clll xl11+s,Tnl¢III_I

+

.v P.+

H_

ottv.,

t

i-,.t °

_rlm |'

f.

.tu 1-

.21_ I

1

t .

0.11-*. " *'. "*" "% "-"

J°"f _ 0.2

0

e+

% "

• :. ":',.

lO

Frsquency (r*_*¢)

PIO ResearchStatusWorkshop @

338
BUFCM-HIGH

q/FLo N CASE COMPARISON

, .._*_ +L]!lltill|It l II lllll l I lllllll ! t[t_,_Z. , __

_a

140

_og_ Tmck_g

_e F,_,..,,¢v h_p

II ° gl° •

• e °l •

"o • •

n ° •

410l I

0 l- .

-HO

I0

1

!':I

1

ou

e I

e°eielem

1 _ • • 11 IIIBi wil|

Frequency (r_d/leCJ

8 April 99 PIO Research Status Workshop

1o

°11o[

BUFCM-HIGH DROGUE

TRACKING TIME HISTORIES

441

o+t rfl't;_ I' i

-44t

o 'Io 20 _Io 40 llo' III

Tune ilK)

1|

t0

o.

.lO

.tl (

ii t I

l , _ i_ I

jP: ' [ ' I : i I ,_i + [.,_" _1 + ; J p

i ' " t ' '1 '(

', + _, t i + ,"

'l(

11me (see)

8 April 99 PIO Research Status Workshop @

339
PILOT INPUT PSD FOR

BUFCM-HIGH DROGUE TRACKING

120

._ 100

m

._ eo

m

_ 60

Q.

e 4O

tu 20

0

o- PIO I Region ]

• PIO 2 Region [

_- No PIO Region J

)

f

• /i _

• ' m

o ,,). . )ore u.

1 2 3 4 5 6 7 8 9 10

Frequency (rad/sec)

8 April 99 PIO ResearchStatusWorkshop

q/FLo N FREQUENCY RESPONSES FOR

BUFCM-HIGH DROGUE TRACKING

8 April99

%C _J

L.

A

le

a.

.560

0.8 t

i.,I 0.4

_0.2[

o

$

_e -i)

10

lO

Frequency (radhmc)

PIO ResearchStatusWorkshop

340
PIO PHASE DELAY

REQUIREMENT

_s_'__!ill I IlL I

8 April 99

o_

02

. 02

t,. t=*¢) o is

Ol

"I

o

o

Hoe*_ (eurc_ow)

o pA-ZO-lO _OR= 10 I=IOR _S

_ 1_ _ O _

F_ - no ID._so

t_e p)o

_._ o

(sAs.oe) He_,,.Up

HO_=S I_.r_so

_loe. 4

I;vP_o

I l I I I t i I

05 1 lS Z 2S 3 3S 4

F**U_,.VStSbmV F,*Won¢V .,__ (_==¢)

PIO Research Status Workshop

CONCLUSIONS

• Frequency domain analysis techniques were

successfully applied to flight test data to obtain

describing functions in the presence of rate

limiting.

• Results display the expected magnitude reduction,

significant additional phase lag, and input

amplitude sensitivity associated with rate limiting.

• Frequency sweeps and drogue tracking runs

allowed for best extraction of PVS characteristics.

8 April 99 PIO Research Status Workshop O

341
CONCLUSIONS

• PIO frequencies and task bandwidths were

identified from the pilot input PSDs.

• Excessive phase delay due to rate limiting led to

PIO for both drogue hook-up and tracking tasks.

• Results from the analysis of the flight test data

support the application of B andwidth/Phase Delay

criteria for the prevention of PIO.

8 April 99 Pro Research Status Workshop

342
COMPARISON OF PIO SEVERITY

FROM FLIGHT AND

SIMULATION

Thomas J. Cord

AFMC/AFRL/VAAD

NASA PIO WORKSHOP

APRIL 1999

PlO FREQUENCY AND

MAGNITUDE

• PILOT CONSISTENCY

- FLIGHT

- SIMULATION

343
10,

5-10 flight _.,_,_._

9i

84

71

6_

5

4

3 ----

2

1

0

L_e --_---" e=

e_q.j

2 4 6 8 10 12 14 16

f0T

5-11 flight _, _._

9_

s_

64.

44

eo

0 2 4 6 8 10 12 14 16

344
10

2-5 flight _,_,-0cy_ 0_.

O

• | n o

o|

e*_q[

_f,,qj

2 4 6 8 10 12 14 16

10

2--8 flight PlO_q_.._._

_ • #O •

m •

0 2 4 6 8 10 12 14 16

345
10

81

71 •

edp •

10 T

t

9_

3-13 flight p_q..=_

I

84

I

s_

I •

41 O •

3_ Oo • • 4)

• oJ

2_

0 2 4 6 IB _0 12 14 16

346
x HAVESIM, PLOT LIMITSA

10 I • x •

X

9_

X X X_'X

4 _ ;, l,

II

1

[

ol

0 2 4 6 8 10 12 14 16 16 20

12 "_................................................................................

HAVE LIMITS

SIM

PLOT B

10 I •

X _

81

e[

X_

,3 t_ X

• &,

2,

•

O.

0 2 4 6 8 10 12 14 16 18 20

347
_2 ..............................................................................

HAVE LIMITS SIM

PLOT D

|

xX X _

2 4 6 8 10 _2 14 16 18 20

PIO FREQUENCY AND

MAGNITUDE

• EFFECT OF SIMULATION ENVIRONMENT

348
10 ,

9 J- 5--9 fli_'_P_lO freuecluen_yssmagnitude

I

84

the po_t _ere _ Ina_ I

__=._-_°_-__ I I

il

,I

11

ol

o 2 4 6 8 10 12 14 16

I IxIkdU_ p#x_t_ abo

7 f ._m tn h_,m .,'uuma,,d

4, ° j4

4. •

• _@I @11

• _:'.-: ,:.• .. ,.

349
5--9 LA ,M_,RS no motion PIO fmquencyversus magnitude

e m_m _'_Jmmced wma.

llnn_n_4Jde _ _her MS- I

orf_. 9

B

7

6 •

4 I •

?,:_ " .,".• . |

• • -- $

2 $

0 .

2 4 6 8 10 12 14 16

rno_n m_oducesless rnagnt_e

5--9 LAMARS notion PIO frequencyversus magnitude --_,_,_-y

9 ! t1, _,_ , ,

sl

et *

s_

4k,

31

21

I(

01

0

sm

4. A

8 10 12 14 16

350
TIME HISTORY

ILLUSTRATIONS

• GROWTH OF PIO MAGNITUDE

• INFLUENCE OF SAFETY PILOT

20

SOS Flight, 2DU with 20 deglsec mtelinit

15

: !!

_o ;,,] _! :,

-_oJ

351
2O .................................................................................................... 0.80

SOS Simulator, 2DU 20 deglsec rate lin_t

15 0.60 I J 10 0.40

" " _ " ' " '_ 0.20

, ',:._V_ ,_:J'_I _ _'_.:,_"_"_i_'-_'_v,_ _ , /o _,

-Io : - " , 4)AO

_ -............................................. 5--10 flight PIO .......................

magnitude,

l chronological

",41

ml

6J

4;

21

Ol

• • • • | •

• • • • x 3( x

0 2 4 6 8 10 12 _4

352
._- ........................................................... 5--10 PIO ..................................................

Z

q

I

magnitude, MS-1

p_B

m

!_i! ̧_ii%!

p_C _D

O_

5 10 _5 20 25 30

18

16

14

12

tO

8

6

P

I

0

........................ 2_.5 PIO-nag, chronological .................................

fliqht

• X

4 6 8 10 t2 14

353
............... 2--5 PIO nag, "

chronological ms-1

x

0 2 4 6 8 10 _2 _4 _5

OTHER OBSERVATIONS

• INFLUENCE OF PREVIOUS RUN

• INFLUENCE OF KNOWLEDGE THAT

TEST IS FOR PIO

354
PIO TRIGGERS

• FLIGHT: NOMINAL TASK PROVIDES

TRIGGER

• SIMULATION: ARTIFICIAL STIMULUS

MAY BE REQUIRED

S UMMARY

• EFFECT OF MOTION - MINIMUM

CHANGE IN RATINGS, NOTICEABLE

IN PHYSICAL CHARACTERISTICS

• SAFETY PILOT - ENDS TASK

SOONER, MAY AFFECT MAGNITUDE

• EVALUATION TASK - KNOWLEDGE

OF PIO TEST MAY INFLUENCE

RESULTS, ARTIFICIAL TRIGGER

SHOULD BE CONSIDERED.

• PIO FREQUENCY - A RANGE NOT A

NUMBER

355
FLYING QUALITIES GROUP

• -1952 Air Force Control Laboratory

• -1962 Air Force Flight Dynamics Lab

• 1979 Air Force Wright Aeronautical Laboratory

• 1989 Wright Research and Development Center

• 1991 Wright Laboratory

• 1998 Air Force Research Laboratory

• 1999 deceased (no FQ research office)

_l. SOSo _r

10 .............................................................................................................................................

8

7

6

5

4

3

2

1

0

0

this ct_l sho_ the * .... A ...........

pibt bac_dngout of Ihe task

for 10d/s. at 40d/s, he is typicaly

got his best b_t d of aggressiveness

md _ for the la,sk, at 157d/s,

he is o_r_ aggressi_

, f f f f

2O 40 60 _0 100

• Hunter

• Taschner

• Hu_el

• Grade

-- Mes,n

MedLan

--rl_x

min

120 140

I

180

356
2d - sos - pior

6 .................. different _ showuphere, hunnel .......

5-4-

4-+-

3-+-

2-+-

0 1

is consistent until he gets to lOd/s, v,here

he backs out of the lask. hunter and glade

stil show g'Bt dip around 50d/s v_ere they

get the best performance v,ithout o_er _' eciting

the system tasc_er has a simbr dip down

at 30-40d/s. A

• Hunter

• Tasohner

• Hunnel

• Glade

-- Mea'l

• .............. -- Meadian

--Max

Min

I I I I

60 80 100 120

I I I I

20 _0 140 160

I

180

2d - dis - chr

10 ............................

sintlar trends to sos task,ecept that _e piiots do not

9 _ I_ck out at low rate Irdts.fudher e_ldenee that discrele

/\ is the best task for pio.

8 '_ / \ this data also sho_ a bt of the pmbbms Wth

, .\/ \ "'+''''°'"

Hunter

2 I -- Median

1 -_ "'_ --max

rRn

0 1 I I I I I I _ I

20 40 60 80 100 120 140 160

I

180

357
2d - dis - pior

6 ............

very si_ to chr

me Hunter

Taschner

I.'tJnnel

Glade

--Mean

m. ,*. A _* _ Meadian

-- ITI_(

.... r_rl

I I t I I , I

20 40 6O 8O 100 120

I I I

140 160 180

358
PHANTOM WORKS

Slab|liB/, Control & Fly_g Quallt_s

A Summary of the Ground Simulation

Comparison Study (GSCS)

For Transport Aircraft

PIO Workshop at NASA-Dryden

April 6-8, 1999

Terry von Klein

Stability, Control, & Flying Qualities Group

Boeing - Phantom Works, Long Beach

PHANTOM WORKS

$11bil_y, ConfroJ & FI)qng Quatifiez

GSCS Goals

Fly a Test Transport Aircraft

- Degraded FCS Configurations

- Evaluate Pilot Induced Oscillation (PIO) Characteristics

Evaluate Identical Configurations in Simulation

- PIO Characteristics

- Motion & Fixed-Base Ground Simulation

• Compare Flight Vs. Simulation

359
PHANTOM WORKS Test Faci I i t i es SlaMIl~. Control 6 Flyhp Oualltbs 

Modern, High Wing 

Transport Test Vehicle 

- Specialized, One-of-aKind Test Aircraft 

- Fly-By-Wire Flight 

Control System - 

- Change-A-Gain (CAG) 

-~~~ 

FLIGHT FCS 

CONDITION CONFIGURATIONS 

High Speed Pitch Phase Lag 

Cruise Condition 

(285 KIAS, Clean Pitch Command 

Wing, 25000 ft.) Sensitivity 

Low Speed Pitch Phase Lag 

Power Approach Pitch Command 

Condition Sensitivity 

(1 45 KIAS, 1 2000 ft, 

Flaps 8 Gear Down) Sensitivity 

Roll Comnand 

PHANTOM WORKS FCS Configurations SWllol,. Conml a FWnp Qwfrucr - ~ 

HANDLING QUALITIES 

EFFECTS 

Add Up to 100 msec of Extra Time Delay in 

Pitch Response 

Increase Pitch Response to Pilot Input By 

a Factor of 2.0 

Add Up to 100 msec of Extra Time Delay in 

Pltch Response 

Increase Pitch Response to Pilot Input By 

a Factor of 2.0 

Increase Roll Response to Pilot Input By a 

Factor of 2.2 
PHANTOM WORKS

StabUl_V, Control & Flying Q_tn_

Pitch/Roll CAG Locations

tAG

pll¢m cornmaced

t4molldllF -- -- "1

¢=I_fIIORI =

PilOt Pltcrl

Jlput

Cammnnd

Iml;m_

Pith aw F4r.Nr4 Elel_l_on

_ p...r¢¢--41 _n'CPollmll_l< I

I

I

I _ _©rm't

-- [lolOrl

P*.J_,G FmOD¢I

PltD Phl|e LI(I

¢_ IOMmllcmu

CJLG

ROll ¢¢1Vlmlt¢l

11iellllllld_ -- -- "I

COnlOIrlll_l I

t

PllOtROll __l_-'] - /_ R¢0I njnPcPNud _ 6Jleforl

1-'-Bin" Polluon rlpUt "_P_ _:_'_ _PI_ Pllth prcce_ln Q JCommalld

¢_mnond

uuma

Wt'crMt

F*,bffOa"i

OOk'J/VO"

PHANTOMWORKS High Speed Evaluation Task

Sfabifi(y, ConUol & Flying Q_slities

• Boom Tracking Behind

Tanker Aircraft

• Separation Distance of

Approximately I Plane

Length

• Pre-Defined Scripts of

Boom Movement

• Feet on the Floor

o

o

%,

<>

n ,606 """ "b

oo_ oo

J lint tt_ t ! '; |

361
PHANTOM WORKS

Steblfl_y, Co4troI • Fly_g Q_,lltlBs

Low Speed Evaluation Task

• Formation Trail Task

Following a Small

Leader Aircraft

• Separation Distance of

Approximately 2 Plane

Lengths

• Pre-Defined Scripts of

Leader Maneuvers

• Occasional Pedal

Usage

/_ XOJ'JAVD"

PHANTOM WORKS

Stabi#ty, C¢_t:o/ & F_ing Oualittes

Testing Summary

• Flight Test

- Two Evaluation Pilots

- One Flight of 5.5 Hours Duration

- Very Few PIOs Noted

- Formation Trail Task Higher Workload Than Boom

Tracking

- Potential for Structural Mode Excitation

• Simulator

- Minimum of Three Evaluation Pilots

- Motion Response

• Valuable at High Speed Test Points

• Of Neutral Value at Low Speed Test Points

- Structural Modes Not Modeled

362
PHANTOM WORKS

Sllblll_y, Conbol & FlYk_l Oualltt_l

GSCS Status

• Very Early in Data Analysis Phase

• Complete Set of Flight Test Data

• Similar Results in Fighter Studies

• Variable Stability Capability of Test Vehicle

- Respect Flight Safety

PHANTOMWORKSGeneral Fit. Vs. Sim. Results

Sfilbih'ty, Control & Flying Qumlil_ez

• Simulator Harder to Fly

- Control of Separation Distance

- Differing Piloting Techniques

- Simulator Generally More PIO-Prone

• Level of Target Aggressiveness

- More Aggressive Target Required in Flight

• Pilot Ratings

- Inconsistent Pilot Rating Trends in Simulator

- More Consistent Pilot Ratings in Flight

• Coupling Between Pitch and Roll Axes

- Degraded Axis Led to Perceived Change in Off-Axis

• Low Speed Motion Cueing

363
PHANTOM WORKS

St=blll_y, Conb'ot • Flying Quatlt_s

Discrepancy Factors

• Simulator Transport Delays

- Visual, Displays of Sensor Information, Motion

• Reduced Simulator Cueing Environment

- Level of Visual Detail

- Depth Perception

- Visual System Field-of-View

- Visual System Alignment to Fuselage

- Motion Responses

• Travel Limitations

• Differing Pilot Input Spectra

- Pilot Adapting to the Situation

- Structural Mode Impact

PHANTOMWORKS GSCS Background

Stebill_y, Conb'o_ & Flying Quatili_

• Sponsored By AFRL/USAF

- Technical Monitors: Wayne Thor & Dave Leggett

• Flight Test Planning

- August 1996 - March 1997

• Simulator Evaluation & Analysis

- April 1997 -August 1997

• Flight Testing

- August 1998

• Data Analysis

- Ongoing

364
365 
“Real (and Imaginary-) Experiences in the Frequency Domain” 

0 Background 

Purpose of Briefing 

0 Frequency Domain 

Analysis 

‘Fundamentals’ 

0 Real Data Analysis 

0 Concluding Remarks 

Realistic Assumptions? 

Not intending to be too “Complex” with this presentation on frequency 

response analyses - therefore, the presentation title is only “Real Experiences in 

Frequency Domain” as opposed to “Real and Imaginary Experiences in 

Frequency Domain.” Pun intended. 

This is the outline of talk. 

What is meant by “Real Data” is experiences where the assumptions needed 

for frequency domain analysis are implicit -- unspoken, but may not be realistic 

or compatible with data from real airplanes. 

In many cases the ease of use of the tools themselves tempt an engineer to 

treat the analysis as a black box. 

366 
0 Purpose: 

Enlighten Users 

(and Analysts) 

Into Practicalities of 

Frequency Domain 

Analyses 

0 Primary Issue: 

Assumptions 

“Engineers Will Typically 

Assume Everything But 

the Responsibility” 

Anonymous Examples 

So the purpose of this presentation is an attempt to enlighten the users and 

analysts involved in fi-equency domain based FQPIO criteria of the errors in 

their ways.. . To champion the cause of common sense over common practice. 

The problem is NOT necessarily the criteria or using the frequency domain - 

the problem is that the analyses for nonlineadreal aircraft data are not trival nor 

are they “independent” of assumptions. The criteria are not explicitly 

considering these assumptions and the users are not aware of the assumptions. 

Engineers are infamous for “assuming” everything but the responsibility . 

Assumptions are always used. Keep knowledge of them and use engineering 

judgment for applying techniques wisely. 

Maybe not such a good idea to bash engineers in front of a roomful of 

engineers. Probably would have gone over better at SETP or at a board 

meeting. Hmmm.. .. 

Anonymous examples are used in this presentation to high1 ight “assumptions” 

- The examples are of using tools, applying these criteria and concepts rigidly. 

The definitions in many cases need revision and clarification. Assumptions may 

be incorporated in the criteria, or distributed to the user, or understood by the 

userianal yst. Wrong answers are being found. 

367 
. General Linear System 

y(t) = AR(w)sin[wt +I$(w)! 

- Partial Fraction Expansion 

- 41 R2I V(S) = - +- + other terms S+JW s-jw - 

The Frequency-Response Function 

of a Linear System 

Is Uniquely Determined By the 

Time Response To Any Known Input 

PlmcU,e, on Comp*mmUv‘ Sol*lon (Transient) (steadystate) 

- For Particular Solution: 

R(w) = IW( s)J. NQ) = arg W( jto) 

Emphasis on FUNdamentals.. . The fundamentals of fi-eq. domain analysis are 

that the response (y(t)) out of an arbitrary system (W) in response to an input, r, 

can be decomposed by partial fraction expansion into essentially three terms 

using Laplacian operators. 

The first two terms are the “particular” solution. The remaining terms are the 

“complementary” solution. 

The “particular” solution is the “steady-state” contribution of the response, y. 

The time response, y, is thus described from the frequency response of black 

box (or transfer function) where R= magnitude and @ = phase of W. 

The key to this fundamental property and why Frequency domain analysis is 

so nice for engineering use, is that “The frequency response function of a linear 

system is uniquely determined by the time response to any known input.” 

The key priniciples/assumptions to remember from this are: “LINEAR’ and 

“Ignoring the Other Terms” 

368 
(o i.p.t = 8.0 radlsec

1 'A

i

_'._

°[ 1

<_ o i _oo,,,(_o.e,o.)

-20

10 -_ 100

°l i

-50 Phase =

- 100 5?'3At_bP_

-150

10' 100 10'

The System, W(s):

]

S 2 4- 2(0.1)(8.0)s + (8.0) 2

"Transient" Behavior

Is Assumed to Be

Inconsequential

Steady-State Yields

One Frequency

Response Point

Frequency (rad/sec)

FllgMRw_ff:hGroup _J'! ViN'ld_r_Eng|notrlffg

• An exampl e of these principles is shown.

• Transfer function of system, W, is as shown.

• Input is 8.0 rad/sec sine wave.

• After transient behavior (assumed to be inconsequential), steady-state can be

used to find phase and gain (freq. response) at the input excitation frequency.

• The opposite principle also works (freq. domain to time domain) since we are

analyzing a LINEAR SYSTEM.

369
250

200 i

-511

o

Transient Behavior

Is Inconsequential

• When Is It Not?

- Prime Aircraft

Example:

Unstable Systems

Time Response

0.5 I ! .5 2 2.5

Time (se(s) I

_uency Response

Bode Plot: 1/(s-2)

°i i i i iiiiii ! i !:::'_:

......!---i-,..... ........

-_°o" io° ld

Frequency (rad/sec)

1 i i J _ i llll J i i t i ill

io" io0 lo'

Frequency (tad/see)

F#ghf Relemrch Group _ _. _,* '. Ver_dl_ln Englnmerlng

• THEORETICAL = do not apply to REAL WORLD

First example of a BAD ASSUMPTION.

• Ignoring "transient behavior"

For example, the best example of when this is a problem is for an unstable

system.

Unstable systems have frequency responses. The uniqueness properties

between time and frequency domain still apply.

The problem is that it is impractical for this identification in the real-world.

From the tim e response, the transient behavior "overwhelm s" the tim e response

and the "steady-state" frequency response characteristic is "hidden" in all

practical sense of the word.

This point will be returned to at a later point in presentation.

370
Why FFTs?

• Extremely Efficient

Algorithms for

Computation of

Spectral (Frequency)

Characteristics

Utilizing Power of 2

Significance in

Fourier

Transformation

• Entire Frequency

Response "Answers"

from One Data Run

Linearly-Varying "Pilot" Input

0 5 10 15 20 25

Time (ec)

.51

0 5 10 15 20 25

Time (ec)

Flight ReseRtch Group _w _,= _,_" Vorfdllln Engineering

• Most practical method for frequency response computation occurs from Fast

Fourier Transformations.

• Extremely efficient algorithm for transformation to frequency domain.

Utilizes power of 2 in time history sample.

• "Entire" answers from one time history.

• Involve a whole set of their own ASSUMPTIONS

371
Time Domain

o:o1 'I,I !tt,tt[K!I

,!,11!;

1 0 10 20 30 40 50

Time (se)

60

o._ A_AA_' I_ L

-0.5 0 10 20 30 40 50 60

Time (se)

Frequenct, Domain

Oi , , , , , ,,,, , ; ; ; ;;;,_

.i0 L ..... 4.--" -- _.--I- _.-_,- !-" 4-F m i ......... : ,'

-20_......... "--i-'-t-,'-,'-, ,- ..... r---, --,--, ,

3 100 10 _

I Frequ=cy (adsec)

0 , . , ; ; ; ;;i;

i ' : , , ,'": : ', : : "::"!

_ 30 _ .................. -_ " F..... ' ' : i i!ii!i_

_-i -9o_-------;---'--'-'-'-"!!- ......... r-T-r_.-::_d

10 10

, o 101

i Frequecy (acfsec)

V_dian Engineering

• Example of time response and frequency response.

• Example showing a "linearly varying" frequency input.

• Note that this is for a linear system.

• Everyone can do them. No pain, no suffering.

• Tools make it easy to apply FFT without looking at the whole picture.

• Of course, now that everyone can do them. Everyone does. Do they all know

the "underlying assumptions" involved in this transformation?

•"Garbage In, Garbage Out"?

• "A Little Knowledge is a Dangerous Thing"?

372
• "Optimal" Input

"Shape" for FFT

Computation?

• Broadband Input?

I Fr_luency (Hz)

U)

Schroeder-Phased Input

-0.0 ;ill,jvVlv;ifilliiJilt l,'l

.1 [

0 2 4 6 8 10

Time (ec)

-401

0 2 4 6 8 10 12

Time (ec)

[

12

F_ht Res_rch Grip _ * _,= _ _ Vertdlan Engineering

• A practical matter, not considered by many, is the importance of the input

exci tation.

• Unlike the "frequency sweep" input, it is not the "optima l/i deal" input

• Schroeder-phased inputs are better. Chirp-z inputs are also better.

• We will visit the importance of input on the next chart.

373
• All (Freq. Sweep) Inputs Are As Good As Any Other

• Considerations:

Input Amplitude / Input Rate / Frequency Content /

Analysis Technique / Fit Condition

-2 0 10 20 30 40 50 60 -2 0 10 20 30 40 50 60

//till

Fltghl Riilirch Group _ _ _ _,* i,, Vllrldllln Enlllnlirlnll

• Another bad assumption illustrated concerning inputs.

• In practical terms, the input for the frequency sweep has to consider: the

amplitude, ampli tude rate, frequency content, analy sis technique that will be

used, and flight condition.

• Agai n, for Singl e-i nput, singl e-output, no noise, lin ear, time-inv ariant system

analysi s, all of these items are immaterial (with exception of frequency content).

This is NOT the real-world.

• Input amplitude: important for signal-to-noise ratio.

• Input rate: important for "rate-limi ting effects"

• Freq. content - determines range of"val id" data

• Anal ysi s technique - ensem blin g of win dowed data usually requires

"broadband" / noise-type excitation across entire tim e history.

Schroeder-phased inputs are tuned to frequency FFT harmonic

frequencies (for lack of a better word).

• MORE DATA -- Better??? Only for certain circumstances

• Flight Conditi on - Tradeoff between "constant" flight condition and accurate

low frequency identification. Phugoid issues in particular. Low frequency

inputs will excite phugoid (i.e., speed changes) - these are "real" effects yet can

be "different" than what some people want (i.e., constant speed approx, for

instance). Have to be careful what you asked for...

374
• Input I and Input 2 Differ Only In Magnitude

_L ...................

F_M R_slmh Group _ "_,. _ * " V_dlan Engining

• An exampl e of input importance.

• System under identification is identical.

• Compari son of two freque ncy re sponses generated usi ng two different siz ed

inputs.

• Very, very different results depending upon input size.

• System was nonlinear.

• Analyst said - "what's going on. You asked for frequency responses and I got

different "answers" every time."

375
r ]

Pilot I --

t Command _ Forward _>_ Path H_ _ Command _ ,;_-_t

I Ra=LI,,,_ I Co,m,, t 1..e_. tl _ t _taLi,,_,IZlt ,_;;_ t] I

,=

Control i

I

-- Digital Flight Control System (DFC$) ........

n

Aircraft

i

i

J

Many Other Nonlinear Elements Abound

Nonlinear Elements Can Be Very

Desirable / Valuable Tools For Excellent

Flying Qualities

F#ghtRwMrchGrOUp _*,,. _.,L: V_-IdlanEnsInaNPrlng

• A schematic diagram of"typical" rate limiter locations. Many other

"nonlinearities" abound - not shown.

• Some limiters are intentional and necessary (ie., the surface command limiter)

- others are phy sic al I imitations (i.e., the actuator)- some are used "erroneously"

(such as the pilot command rate limiter) because HQDT "requires" it. (For

instance, if max. value, unrealistic inputs are used just for "PIO" evaluation, an

easy solution for the designer is to slap a "pilot command rate limiter" in the

forward path. The result is that a "PIO" will not happen for the unrealistic

HQDT task. However, the real result is that 20-25 msec of time delay is now

added to the flight control system and the potential for a real PIO is increased

just because some people teach the wrong thing for HQDT.)

• Nonlinearities are not bad. In fact, they are quite the opposite. They are

necessary for good FQ. The only problem is making sure that the FQ tools can

identify these "good" qualities and not legi slate against them.

376
This is not an LTI system 

Issues in Frequency 

Response Derivation: 

rn Single-Input, Single-Output 

Linearity 

Time-Invariance 

- Stationarity 

Uns fa fed Ass u m p t io n : 

Linear Time-Invariance (LTI) 

THEORETICAL basis = do not apply to REAL WORLD 

The assumptions in fieq. response derivations are: 

(Many times, but not necessarily) Single-input, single-output (Le., output is 

caused only by the one input) 

(Always) Linearity (ie., linear system is q=M,a+ ..., nonlinear system is 

q=M,,a2 etc. ) 

(Always) Time-invariance (ie., y = function of time) (Stationarity is the 

“controls engineers” term for time invariance) 

Linearity conditions are easily violated by changes in flight condition, position 

and rate limits, breakout force, friction, hysteresis, nonlinear command 

gradients, etc ... . 

Time variation is also a rate limiting effect. In other words, the FFT analysis 

is assuming that over the time period for the identification, that the system has 

not changed. 

377 
Time Domain Frequency Domain 

I 5 

f 

I 

I 

I 

OWPD~D~WLDP 

! f, 

01 1 W m 

Fnrm6 # 

OWPDODWWWP 

la 

kbd 

Can rate limiting affects be identified in Freq. Domain? Yes. Here’s an 

example. 

Note phase rolloff and amplitude attenuation. 

However, the most important condition for this result is that the rate limiter is 

no longer “time varying” - it’s a quasi-steady. See rate signal above. 

HOWEVER, hard part - for this to occur, amplitude and frequency of the input 

to the rate limiter element depend on lots and lots of factors in real situations 

that cannot typically be predicted or repeatable from run-term, pilot-to-pilot, 

etc. 

Particularly for rate limiters that are “buried” in a control law - that is, the 

inputs depend not only on the pilot inputs but also on the feedbacks, etc. A 

prime example is the actuator command rate limiter shown on a previous slide. 

37s 
0

0

!

0

Time Domain Frequency Domain

l[]fl Illl

]1]11 III] I'f 

1

Flight R_UlCh Group _ _,. _,_ _ V_'ldlan Englrl_lorlng

• Here's a more "typical" example. Note variation in rate limit. Also, noise is

added to input and output. (Not a laboratory condition!!)

• Introduce "coherence function" at this point.

Purpose: Evaluation of"goodness" of FFT.

Real name: "Ordinary" coherence function for SISO case.

• Coherence lets analyst know if FFT/freq. resp. is "valid"

• Not valid (ie., coherence values go <1) if:

1) Extraneous NOISE is present in the measurements

2) System relating x and y (input and output) are not linear

3) Output is due to input as well as other inputs -- not SISO

379
• Ignore Significance of

Coherence

• "Ordinary" Coherence <1.0

Noise

Nonlinearities

Not SISO

• Coherence "Significance"

Has Been "Lost"

• System Identification From

Tracking (SIFT)

AFFTC-TR-77-27, Nov.

1977, Twisdale & Ashurst

• Must Re-Establish Its Role

Flight Resemrch Group _. , ,= _ i_ Verldlan EnBIne_rlng

• Reiterate: Ordinary Coherence < 1 - Noise, Nonlinearity, Not Single-Input,

Single Output (ie., multiple inputs, turbulence, etc can cause violation of SISO)

• Can't just "ignore" coherence - have to understand why coherence does equal

1.0. Involves more analysis of the input and output, and tracking the error.

• Coherence has been used as a "discrete" ie., if coherence>0.6 data is "good"

Not a good thing to do unless you make that level very stringent (coh>0.9,

>0.95). Can be dangerous (Bad Assumption). Coherence is similar to

correlation coefficient analogy. 1.0 correlation is "perfect." Correlation = 0.6,

correlation to real data is not good. Many examples of coherence >0.6, <0.9

where data was "bad." (i.e., not what was expected. If left un-investigated,

would have gotten wrong answer)

• More appropriately, coherence is directly relatable to error in frequency

response estimate. This significance has been lost! (Twisdale did this 20 years

ago!)

• Must get back to its signficance if frequency response analysis is going to do

anything for us.

•Answers from criteria using this data will tend to be regions rather than points

380
• Following

"Established" Rules

• Equivalent Systems:

Typical Range for Match:

0.1 to 10-20 radlsec

- Ignoring Coherence,

or

- Using All Data Points,

Thus, Distorting

Weighting Functions,

or

Identification /

Inclusion of Low and/or

High Frequency LOS

Terms Beyond "Valid"

Data

VISTA " Data TS69f04 Configuration A

_-_,o, _o 1o, ,o,

I

200 I

i2 F......... : :-; i_ ,I'....... ' . :: ,f ........ : : :::::_

.

bO' tO=

1o" _O_ I

• I I |_ ......

105 .... ' ' ' .....

10" ' to ° ,0' 102

I Frequency (rld/la_:)

I I

Ii I ,"alidDataLll

FNght Reswrch Group _m_ _. _ i" Vertdian EnginNrlng

• We've had experience where - after "derivation" of a frequency response, the

"rules" are blindly followed for such things as an equivalent system.

• Neglects phugoid, high order & nonlinear dynamics, structural dynamics,

sensor dynamics, and recording filters. Assumes constant flight conditions.

• Coherence has been ignored (see previous slide)

• Persons have used "all the data points" from a FFT for equivalent system

derivation. Thi s inappropriately wei ghts the hi gh frequency equivalent system s

match at the expense of the low frequency due to the 1/dt frequency spacing of

the data (more pts at high freq., fewer at low freq.)

• Although the freq. range of valid data was "narrow," extrapolation outside the

range was allowed to get a "equivalent match." Unfortunately, answers can b e

MISLEADING.

381
Configurations

2D: stable with

rate limit in

command

path only

2DU: unstable

augmented

to get 2D

characteristics with rate limit in feedback

Flew with rate limits from 60 to 10 deg/sec

Fltgt_l R_el_ch Group _ _ _ I" Veridiarl Engln_l_rlllg

• This is a Simulink diagram used for the "Have Limits" flight test program.

• This model was used to assist the engineers in vi sualizi ng the set-up of the

experiment.

• Subsequent to the experiment, this model has been distributed to users to aid in

anal yzi ng the "Have Limits" data.

• Key "feature" in the data base, analysis, and set-up for the "Have Limits"

flight test is Configurations 2D and 2DU.

• Config 2D has the rate limiter in the forward path only.

• Config 2DU was a simulated unstable airframe - using analog feedbacks,

without rate limiting around the NT-33 Airframe - with an outer loop feedback

structure to augment the simul ated unstabl e airframe to match Confi g 2D

dynamics. The key difference is that the rate limiting term includes the

feedbacks for Config 2DU and an unstable airframe.

382
Problem With 2DU

is Instability

• Rate Limiting

Caused "LockOut" of Control

• With 60 deg/sec

Rate Limit

CHR: Two 10's

• Same Rate Limit

in Forward Path

Showed No

Noticable FQ

Degradation

Configuration 2du; Rate limit = gO deg/sec; FFT Method

i ,. f r . I [ 1{o2dIsraeg/le<l/

p I I I_l[ll_

IE

°! !lllll I

ol I t l, llt! I l ilE! ]! /

Frequency (rad/sec)

Flight Rm_rch Group _ _ _,. ;_ V41_dl&n Eng|nl_wlng

• In a very brief summary, a key conclusion from the Have Limits program is

that Config 2DU have very poor flying qualities. Pilot Ratings were 10 for the

least amount of finite rate limiting (ie., with 157 deg/sec rate li mi ring -

essentially no rate limiting, 2DU got ratings of 2, 5, and 4. But for as little as 60

deg/sec rate limiting, two 10's were given.

• The FQ deficiency for Config 2DU was loss-of-control. Once the aircraft was

on the rate limit, the feedbacks were locked-out and the aircraft entered a

departure scenario. (NT-33 VSS was disengaged upon loss-of-control).

• Same rate limit, in forward path, was not a noticeable flying qualities

influence.

• Using the Simulink model and assuming a pilot inputsize, "rate limiting"

effect in frequency domain is noted.

• Issues:

1 - have to "assume" a pilot input size;

2 - can't get freq. domain "answers" for rate limit values < 90 deg/sec

Only done analytically, not flown.

383
With Control

Lock-out Due to

Rate Limiting

• Incoherence

Time-Varying

System

• Identification

of Unstable

Aircraft

Without

Stabilization

/ I t

_-C°nftguraU°n 2d u; Rate limit= 20 deg/sec; FFT Method = tfcomp2; SWp = chirp I

ol _i I iTlkl _i [Io2. 15Zdw/_._1 ' i +2du 20dea/sec[[

1

10-_ I_ _ _-_

Frequency (rad/sec)

!

FUgM Reseerch Group _ _,. _,* =,

• As example, for 20 deg/sec rate limit, the frequency response data for 2DU is

garbage. Reason: the aircraft hits a loss-of-control issue. Time varying system

with nonlinearity. Also, once aircraft is in rate limiting, the feedback is

"ignored" and the bare airframe ch aracteristics are what is being identified.

• The results are essentially not valid.

384
Time Response shows

when the rate limit is

encountered, 2DU

reverts to unstable

open loop system.

FFT-derived frequency

response is not valid

Control Of Vehicle "Lost" - Departure

_ Is0O_,, Config _. , ; Rate limil = 20 deg/sec; Swp = chirp

-2000 20 40 60 80 100 120 140 160 tB0 200

Time (sec)

Fltgh_ R_'_h Group _ * _,. _* :R Verldlan EnglmNwln g

• Here the time history really shows what's going on. Specifically, like the

earlier exampl e, the transient response is NOT negligible.

• Once aircraft is in rate limiting, the feedback is "ignored" and the simulated

unstable bare airframe characteristics are driving the response

• Once the rate limiting starts with Config 2DU loss-of-control occurs. Note

the time histories where alpha goes +/- 25 degrees and the g's go way beyond

+/-2 g's. (The plot is artificially limited to +/- 2 g's)

• FFT-derived frequency response is not valid since it is no longer linear

aerodynamics or tim e invariant.

• In fact the response immediately goes beyond the scope of the small

perturbation model.

• These agree with the results experienced in the flight experiment.

385
0 Frequency Response 

rn Extremely Valuable 

rn Most ‘CommonDerivations 

Information 

Knowledge’ Properties 

Only Pertain to Linear 

System Analysis 

rn Caution I Care Must Be 

Used In Real Situations 

Particularly Nonlinear, 

Time-Varying Systems 

Analysis 

- Le., Today’s Aircraft!! 

Said enough. Just summarizing the points ... 

9 Don’t let them kill the messenger, Andy. 

Reiterate that Freq. Domain analysis IS a powerful tool - very useful. 

However, it can’t be used carelessly. Unfortunately, it is ... 

I’ve cited some examples. Many, many more were available but I couldn’t put 

them into a 30 min. presentation. 

386 
0 Tools & Techniques for 

Proper Analysis Are 

Ava i lab le 

e.g., System 

lndentifcation From 

Tracking (SIFT) 

0 Retain Engineering 

Judgment in Analyses 

0 Scrutinize Assumptions 

0 Develop ‘Standards’ 

Reiterate that tools are available or can be developed. Not rocket science. 

Clearly, evidence abounds that the fundamentals of frequency domain analysis 

are being ignored, forgotten, whatever - but things will get worse if they don’t 

stop, step back, and think about what is being proposed and done. 

Standards for analysis will help. 

387 
• Criterion Indicates

"PIO" Problem

• AIAA-99-0639

"Determining

Bandwidth in the

Presence of

Nonlinearities"

• FQ Data Shows

Loss-of-Control for

Config 2DU

• Correctly Predicts Pilot

Rating for Wrong

Reason?

ol plop_ _P,o _'_s7

: ,.-=.., y

Re(: AIAA-99.0639

FJght Ru_rCh Gr_p _= _,* :, Vorldlan Englneerlng

• In AIAA paper 99-0639, frequency domain data was presented for these cases.

• Don't know how these data were generated - can't repeat analysis.

• Further, they should show unstable aircraft behavior. They don't

• Finally, the frequency responses in 99-0639, show a feedforward, time delay

effect of rate limiting - not the loss-of-control issue. That's what the bandwidth

criteria, shown on the plot, indicate.

• Basically the criteria are predicting the right answer for the pilot rating, but for

the wrong reason. The real data - the pilot comments - don't match the criteria.

The criteria doesn't say "loss-of-control" for this configuration.

388
Simulink Model

• Uses Small

Perturbation Linear

Aircraft Model

• Not Intended for

"Nonlinear" PIO

Analysis

- Used for

Visualization of

Aircraft Set-Up

- Small Perturbation

Checkcases

Control Of Vehicle "Lost" - Departure

/

Configura/,'l 2du ; Rate limit = 20 deg/sec; Swp = chirp

-EI 0

"time (sec)

Fliglff Resw_rch Group ,a_w v= _,* _ V4N'idian EngllNN_rlrlg

• Another problem with these analyses i s the use of the Simulink model.

• The model was intended for visualization by Calspan and AFTPS engineers of

the experiment. It was also used for small perturbation checkcases.

• The model uses a simple three degree-of-freedom, small perturbation math

model of the NT-33.

• The scope of the validity of this model has NOT been determined. However,

clearly, it is not valid once the rate limiting occurs with Config 2DU and lossof-control occurs. Not the time histories where alpha goes +/- 25 degrees and

the g's go way beyond +/-2 g's. (The plot is artificially limited to +/- 2 g's)

• Again, the model was never intended for the purposes that it may be being

used for at this time. This should have been obvious from inspection of the

"aircraft" model form.

389
q
Pilot Modeling for Resolving

Opinion Rating Discrepancies

David B. Doman

Air Force Research Laboratory

April 8, 1999

• Inter/Intra pilot opinion rating variability has confounded flying

qualities engineers since the inception of the rating scales

•A method for extracting quantitative information from

experimental data to provide insight into rating variability and help

gauge the validity of ratings would result in a valuable engineering

tool.

•Idea #1 Extract metrics developed for pilot-in-the-loop flying

qualities criteria from experimental frequency response data.

•Idea #2 Estimate a range of ratings by using highly accurate

models of pilots and varying physiological parameters over a

reasonable set of values.

391
PitchAlfi[nde

Command P2ch Error Aircra_ PitchAttitude

I " PilotFunchon Describing Air_rat_ Dynaxr_cs

Neal-Smith, Bacon-Schmidt, Efremov MAI:

•Closed-loop resonance

•Pilot phase compensation, (Pilot phase excluding

neuromotor lag and time delay)

•Each assumed all pilots behave the same

Neuromotor lag (related to aggressiveness) and time delay

vary over pilot population, What ran_r___ge of pilot ratings can

be expected?

392
!i/?:!_ili!i ii:!i_ii!Ji iii! ̧:

Assumptions

• Compensatory Tracking (SOS)

•Minimize mean squared frequency weighted tracking error

subject to human operator limitations

J = E_(e_ + fu_)

--_Tks+ le(s) ef(s)

7_s + 1

Control rate weightingfdirectly linked to pilot's neuromotor

dynamics.

20 Pea_ l".

®-_"o._=-!, .... i---!--i--: :',/: _ !ii!!i!__: i i ii ""+:-+" i i ! ! !'! !ii -"+"+'""! _i!i__.,.,._i ! ! !ii

_°ti/ii::-[iiii ....iUiiiiii ....7-:,i!r??:,

10_ / 10° 10_ "r" ,'r, 102

Phase Dr_op_ow Frequency Lag:Lead - 'k,lk ......

s.2oo t .... .__.,,....,,_.,,_- !!._--:---:--.-:- !!_,_

o 1 ', ', ',III',', 1 : ', ',i',II', 1 i ', ',''

....

_-_ooI....!---_--_-i-___1-1

104 100 101 102

Frequency (rad/sec)

Tp

393
) i/ ̧ i [ : _i i?!

B a co n.S ch m idt and N S-2D

12 , , ,

10 "_

Level 3

\,

8 r, -0.3s,_ --*'I r, -_

_,=o2,o/---,- _ L_,,_I \

/ J_ N8ttttntl Bacon-Scl'_t 6t Pot_ fe,_S2D

-50 0 50 100 150

Lag-Lead -------+

Equivalent Pilot Phase C ompensal;ion

Evaluation of NS-2D (USAFILAMARS)

Full )rder Pilot Model for Neal Smith Config. 2-D Tracking Error Loop

2O

,',. : ', : ',:',',: ', : ', ', ::',',: ', ', ', ', ',',',',1blot N

_ ,,,,_::,,.,., ,, : .,,,::;_:,.. ,, "_r.-1/8

rn.o= 0 __, ' ' " .......'" "--'""_;;_ ' '" _"_'-4t_'_.'-,__"_'_"_"' ' ' ' ' ii1

(_9 " ,, ,, ,, ,,,,:,,',,, : ,, ', ,, ,,,,:,,,, ,_'_ruot &

.... .---,--:-,-_ .... ----.---.--:-,-_ .... -.... ,---,-_-,- ... _. = 1/6

10 4 10 0 10' 10 2

s o,.... ._--'--'$-!i_-'---'--_-;_,"i .... "---i--_-i-_!i-

_> ', ',',',.',',',", _ ',_',_,_',___.:_,'Z4_.', _',',_',',',',',', _-200_.... _---,'-,--, r_ ,r .... ,- -,- ,-,-v,_,_6--,---, "_-,-_v

', ,, ', ',',::',', ', ', ', ',',:",', ', ,, ,, :',,,,,,

_-4 0o_....:--v-:vr:':r.... v-v-rv::::_-:-_--,'-r-,"::_:

_,oot.... .....

104 10o 10* 102

Frequency (rad/sec)

394
"t'.=1/13

20 ................... _ ........ i max

10,- ....:---:-'-'-':::: ..... "--:-'-'-'!i_i -_--:-_-_-:::,. 1/8.5

i ! !iiiiii i ! !if::: : :::::: "t" =

........... oo_'t: iiiiiiii!i!i! : ::::' iiiiii! ' :::'" max=76

-o

.101 ....

10 1 100 101 102

_=_0_- .i.!.:_:ii::.... i--i-'-_-':_!ii .....iiiiiiiiii iliii

"_-200 .........

a. -400 ...........

-600 ............. "_ ......

101 100 101 102

Frequency (rad/sec)

, Conclusions

•OCM methods have the potential to describe differences in and

among pilots in closed loop compensatory tracking tasks for

linear controlled elements.

•High frequency roll-off characteristics of the human appear to

be higher than 1st order as predicted by OCM.

•Performance and workload metrics extracted from OCM fits to

experimental data could provide insight into rating variability

and possibly help gauge the validity of ratings.

•Use as a predictive tool to estimate the range of ratings that

could be expected from a pilot population by varying time

delays and neuromotor lag time constants over a reasonable

range.

395
q
- Acknowledgements

- Closing remark

i _.,.hk_i>_<.,'

• Paul Steinmetz

• Frank Newton

. Everlyn Cmciani

° Patti Pearson

• Dennis Calaba

• Darlene Homiak

, Dave Mitchell

• Moderators

• Presenters

• Attendees

• Tour organizers

• Ed Schneider

397
"All happy families are alike, but each

unhappy family is unhappy in its own way."

Leo Tolstoy, Anna Karanina.

o "All good aircraft are alike, but each bad

aircraft is bad in its own way." Mary Shafer

398
Appendix 1

399
q
Pilot-Induced Oscillation Research:

The Status at the End of the Century

NASA Dryden Flight Research Center

Edwards, CA

6-8 April 1999

For well over a century, as long as people have been gliding and flying, aviation safety

has been threatened by pilot-induced oscillations (PIOs). As, our calendars prepare us

for 2000, the time for reviewing the status of PIO research is at hand. NASA Dryden

Flight Research Center is pleased to sponsor an open workshop doing just this in a threeday session on 6-8 April 1999.

The last public presentation of PIO research was in 1995 and since then, a number of

major PIO research programs have been completed. The results of these programs will

be presented at this workshop, as will be the results of other studies, hypotheses, and

proposals for further research.

The only restriction is that discussion be limited to safety-related PIO; possible topics

include criteria, simulation and flight testing, the pilot's role, design considerations, recent

experiences, rate limiting effects and minimization techniques, civil certification, military

acceptance testing, analytic techniques, and more. In no way is this the entire list of

possible topics and your participation, discussing any topic you feel is relevant, is

solicited. It may be that the coffee-break talk alone can offer some insight into a difficult

problem you have.

As this is a workshop, with short notice, the expectation is that presentations will not be

as formal as conference papers. Copies of the presented material, with whatever

supporting material the presenter offers, will be produced. If possible, the entire

workshop will be videotaped and copies will be available.

This workshop will be unclassified and open to anyone interested, regardless of affiliation

or citizenship. There is no fee for attending. For planning purposes, however, an

estimated attendance is required; the response form indicates a variety of methods for

responding, however tentatively. Requests to attend must be received by 19 March.

Presentations must be proposed by 5 March. Presentation requirements, as indicated on

the response form, must be received by 19 March. Dryden can support viewgraphs,

35mm slides, videotape, and PowerPoint projection (other software requires providing

PC-based software). Advance submission of presentation material and supporting

material will aid the production of copies for attendees before the end of the workshop.

Presentations are nominally scheduled to last 30 minutes, with 10 minutes for questions.

Should this be insufficient, please explain the need for more time on the response form.

Please circulate this announcement to anyone you think will be interested. Anyone

interested in handling qualities, PIO, aviation safety, pilot-vehicle interfaces, and related

topics should be informed of this workshop, as other forums for discussing such topics are

no longer common.

Please respond quickly if you think you might attend,

particularly if you are considering making a presentation

401
Pilot-Induced Oscillation Research:

The Status at the End of the Century

NASA Dryden Flight Research

Edwards, CA

6-8 April 1999

Center

Attendance (Reply by 19 March, please):

Your full name:

Name you want to be called by, for badge

Affiliation

Address for further

mailings about

the workshop

Telephone

E-Mail address

Preferred method for further contact: __ Mail

Presentation (Reply by 5 March, please):

Title

Fax number

__ E-Mail Fax Telephone

Co-Authors

Presentation media:

Special requirements

Viewgraph

PowerPoint

35mm slides

Other software

Videotape

Other medium

Send this form, as soon as possible, to:

NASA Dryden Flight Research Center

Ms Mary Shafer

Mailstop 4840D

P.O. Box 273

Edwards, CA 93523-0273

(805) 258-3396 (workshop only) or (805) 258-3735 (regular number)

(805) 258-2586 (Fax) or email to Mary.Shafer@dfrc.nasa.gov

4O2
Pilot-Induced Oscillation Research:

The Status at the End of the Century

NASA Dryden Flight Research Center

Edwards, CA

6-8 April 1999

Presentations Information:

All speakers who prepared their presentations with PowerPoint are implored to bring a copy on disk, plus

a duplicate disk, for direct projection. We will have the projector and a computer with the software and

would greatly prefer to project the computer version rather than resort to using transparencies. We find

that the projected computer image is superior to the projected viewgraph. Speakers who used other

software can also project directly if they can bring a laptop or a version of the software that allows reading

the images, although such speakers would be wise to bring viewgraphs as a backup on the off chance

that this won't work. E-mail me if you didn't use Word or PowerPoint and we'll see what we can do.

Speakers who are using the projection system are asked to bring a paper copy for adding to the

handouts; if color is important to understanding the viewgraph, I can make a limited number of color

copies, I think.

Any speakers who want more than 30 minutes for their presentations should let me know immediately.

More time is available, but I can't allocate it unless I know who needs it.

The preliminary schedule has, as is inevitable, changed, but most of the changes are to the order of

presentations within session. Speakers whose presentations have been moved to other sessions have

been consulted before the move was made. VIIsend out a revised copy by Friday.

SR-71 Tour:

I'm still working on getting permission to have the SR-71 tour. If it is granted, the tour will be during the

second half of the time set for lunch on either Wednesday or Thursday and the schedule adjusted

accordingly on the other day. For those not familiar with hangar visits, there are just a few obvious rules.

1. Stay 15 ft (5 m) back from the aircraft unless the crew chief gives permission to come closer.

2. Don't touch the airplane without permission

3. Photos are allowed, but flash bulbs (not built-inflashes, but the actual bulbs) are not allowed

4. If we are allowed to look at the cockpit, secure all loose items in shirt and jacket pockets, so that they

don't fall into the cockpit and FOD it.

5. Watch your step, as there are cables and hoses on the hangar floor.

Getting Here

For those flying into the Los Angeles area, it will be necessary to drive to Lancaster (where the hotels are)

and to Edwards. There are a number of airports in the area but Los Angeles International (LAX) is the

most likely destination, although those who can fly into Burbank will find the drive shorter and easier. If

you're arriving at LAX, you will take Century BIvd to the San Diego freeway, the 405, and get on it going

north (Sacramento is likely to be mentioned) by going under the freeway and then right onto the on-ramp.

Go north untilthe 405 merges with the Golden State freeway, the 5, and keep going north (this is the

easy and obvious thing to do). A few miles beyond that take the Antelope Valley freeway, Hwy 14, north.

This splits offthe 5 on the right side and the city name is Lancaster. Stay on Hwy 14 until you get to

Lancaster and then follow the instructions below if you're going to your hotel.

403
If you're arriving at Burbank, turn left out of the airport and go to the Hollywood Freeway, about two miles.

Get on it going north and when you reach the 5, get on it going north. Keep going until you get to Hwy 14

and then proceed as described above.

To get to Dryden, take Hwy 14 north to Rosamond and exit at Rosamond Blvd, going east, to the right.

Stay on Rosamond Blvd. In about 10 mi, you'll come to the Edwards AFB guard post, where you must

show identification. Those of you with DOD or NASA ID will be waved in when you show it to the guard.

Those with other forms of ID should do as directed by the guards. Pre-registered attendees will be on a

list for admission. If there's any difficulty, tell the Air Force guard that you're attending the NASA PIO

Workshop; if there's any further difficulty, ask the guard to call 258-3273

Dryden is about 10 mi beyond the guard post; stay on Rosamond Blvd though Main Base. The road will

narrow to two lanes (from four) and you may think you've gone too far. About a mile after the road

narrows, you'll see a number of metal bleachers on the left. The road to Dryden is on the right, just

beyond these. There are signs, of course, and you can see Dryden down on the lakeshore. Turn right,

cross the railroad tracks, and turn right at the second opportunity, just before the HL-10 lifting body on a

plinth. Turn left into the parking lot right after you go by the F-104G, X-29, and two F8s. Walk to Visitor

Registration, just across the street from the X-15 mockup, and go to the workshop registration desk.

Amenities:

The room we're meeting in is adjacent to the cafeteria. It is open for breakfast and lunch and also for

breaks. The afternoon breaks will begin before the cafeteria closes at 1400.

The Dryden Museum and Gift Shop is in the same building and is open to the public. The Gift Shop sells

film in addition to a variety of aviation and space-related souvenirs, including tee shirts, models, toys,

pins, photos, and similar goods. They now take credit cards.

The Dryden Exchange, inside the facility, sells stamps and common over-the-counter remedies and

toiletries (the cafeteria sells some remedies, too); access is easily arranged. The Dryden credit union can

handle minor financial transactions, such as cashing traveler's checks (in US dollars); again, access can

be arranged.

Dryden has public tours twice a workday; anyone willing to miss a portion of a session can go on the tour

if there's enough space. Additionally, AFFTC runs a tour of Edwards on Friday morning, so anyone with

an extra day can do the AFFTC tour on Friday morning and the Dryden tour on Friday afternoon. Let me

know if you want to do this, as reservations are required.

Lodging:

The better hotels are in Lancaster, which is 35 mi (and about 45 minutes, counting parking) from Dryden.

This list is just a few of them, mostly with restaurants and all the usual facilities. Members of the AAA can

find a more complete list in the guidebook for California.

Desert Inn

44219 Sierra Hwy,

Lancaster

661 942-8401

661 942-8950 fax

mkt_dcscrt-inn.com

Government rate $60 + tax, corporate rate $62 + tax

Leave 14 at Ave K, turning right (east), go a little over a mile to Sierra Highway (just before the railroad

tracks) and turn left. The Desert Inn is a little more than half a mile, on the left.

404
Antelope Valley Inn

44055 Sierra Hwy

Lancaster

661 948-4651 (800 528-1234 for Best Western reservations in US)

661 948-4651 fax

Government rate $63 (includes breakfast & 2 bar drinks every day), corporate rate $63 + tax

Leave 14 at Ave K, turning right (east), go a little over a mile to Sierra Highway (just before the railroad

tracks) and turn left. The Antelope Valley Inn is about half a mile, on the left.

Inn of Lancaster

44131 Sierra Hwy

Lancaster

661 945-8771

661 948-3355 fax

Government & corporate rate $58.85 (includes breakfast every day, dinner Tuesday and Wednesday)

Leave 14 at Ave K, turning right (east), go a little over a mile to Sierra Highway (just before the railroad

tracks) and turn left. The Inn of Lancaster is about half a mile, on the left.

Oxford Inn

1651 West Avenue K

Lancaster

661 522-3050 (800 522-3050 for reservations in US)

661 949-0896 Fax

Government & corporate rate $55 + tax (Continental breakfast and happy hour included)

Marie Callender's Restaurant on premises

Leave 14 at Ave K, turning left (west), going under freeway. The Oxford Inn is on the right, quite close.

The Essex House

44916 10 thSt. West

Lancaster

661 948-0961

661 945-3821

cssexhouse_/_hughes.nct

Government & corporate rate $62 standard room, $74 king, $78 suite (Buffet breakfast weekdays,

continental breakfast weekends)

Leave 14 at Ave I, turning right (east) and go a little over a mile to 10th Street West, turning right. The

Essex House is about 0.25 mi, on the left.

405
One loose end to tack down and some information on the local climate for people not

familiar with the Southern California High Desert.

For larger PowerPoint presentations that won't fit on a diskette, there are two other

options, CD-ROM or Zip. The laptop we'll be using for projecting has both a Zip drive and

a CD-ROM (DVD, actually) drive.

Weather and what to wear:

Dryden is an informal place and I suggest that attendees adapt to the local standards.

Business/government casual, which for engineers starts here at jeans and tee shirts and

goes on to a point just short of dress shirts and ties (and for pilots starts and stops at

flight suits), is suggested. I'm sure everyone will reach a proper balance of comfort,

casualness, and appropriateness. As it is Spring here, a layered approach is often wisest.

The average high temperature for the week of the workshop is 70 degF (21 degC, if I've

done the conversion correctly) and the average low is 42 deg F (5.6 degC). The average

precipitation for the entire month of April is 0.01 in. (0.3 mm), so we're unlikely to have

more than a trace of rain. I personally expect clear blue skies for the entire workshop.

However, there is a fair chance of some wind, in which case the highs will be lower and

the lows will be higher and, more to the point, the so-called wind chill factor will make it

seem even colder. Right now, on Wednesday, 31 March, we've got a cut-off low in the

area and it's blowing about 30 kt, maybe a little more, and the temperature is about 55

degF (13 degC), so I've got a lined jacket instead of the shell I use to keep off the

morning chill.

We'll either have lovely spring days with blue skies and comfortable temperatures or we'll

have windy, cool spring days or a combination of the two. This is why I suggest layers--a

short-sleeved shirt with a wind-proof light jacket over light to medium-weight slacks or

trousers. Just in case I've been overly optimistic about the rain, an umbrella might not be

a bad idea. However, even at its worst, the weather shouldn't be terrible, just a bit

uncomfortable. It is Spring, a freeze is unlikely, and trees and bulbs are flowering. There

may even be some wild flowers to see, although we didn't get enough rain in the winter

to make a big show and it's too early for the California poppies.

406
Attached in MS Excel format is the almost-final version of the schedule (agenda). If you

can't read this, there's a version with CSV comma-delimited text (agendatxt), although I'm

skeptical about its readability. Flat text doesn't seem to be an option.

However, it probably doesn't much matter, as long as you show up at 0800 or so on

Tuesday. Everyone getting this e-mail will be on the list for the USAF guards to admit, so

there shouldn't be a problem.

I'm looking forward to seeing everyone and I think we're going to have a good time.

We will be allowed to see the SR-71s; I'm now negotiating whether we will be allowed to

look inside the cockpit.

Tom Cord is arranging a social event at the Officers' Club (Club Muroc), probably on

Tuesday evening. It's not an official event, but attendance is encouraged.

The Weather Channel is currently predicting "cool" temperatures and rain showers on

Tuesday, moving out on Wednesday, and warmer on Thursday. This is coming down out

of the Gulf of Alaska and may miss us, but probably won't since I've gathered so many

people together here. I interpret "cool" as around 50 degF, by the way.

Regards,

Mary

PS. Ifanything despe_te requi_s you tocon_ct me overthe weekend, you may call me

at 661 942-7434. MFS

407
To: Members of RC Branch

There will be a workshop "Pilot-Induced Oscillation Research: Status at the End of the

Century" here at Dryden on 6-8 April. I have attached the almost-final agenda (in Excel).

Pat thinks it important that members of the branch participate as much as possible in this

and I'd like to invite everyone to stop by for as many presentations and discussion as you

can manage. The people speaking and attending are all well known and highly regarded,

so we'll have a chance to hear the latest news from the people who really know.

Nothing special is required for Dryden personnel to attend. None of the material

presented is classified or limited in distribution. I will have copies of the material

presented for those who can't make it, although the discussion is often more interesting

and informative than the actual presentations.

I hope to see many of you there.

Mary

408
409 
i

410
Appendix 2

411
q
This presentation gives an overview about results of PIO-investigations 

obtained from a flight test program on DLR's flying simulator AlTAS 

(Advanced Technologies Testing Aircraft System). AlTAS is a small civil 

ak, which has been developed as a full Fly by Wire In-Flight-Simulator 

with a safety pilot in the right seat. 

(This presentation has been prepared by Dr. Holger Duda and Gunnar Duus 

and myself) 

413 
I 

o Aircraft-Pilot Coupling 

o Prediction of APC 

o The OLOP Criterion 

o Recent Flight Test Experiments with AlTAS 

o Data Analysis Techniques for APC Assessment 

1 o Conclusions 

l o Future Activities 

. I 

PI0 Workshop, NASA DrvrlPn Riaht Research Center Edwards. Ck 6-8 Am1 1999 

The contents: 

- 1. The aircraft-pilot coupling phenomenon is illustrated briefly. Criteria for 

AFT-prediction are discussed, emphasizing the OLOP-criteria for 

prediction of nonlinear APC. 

-Thereafter the main results of recent ATTAS-experiments, with respect to 

experiment-design, results and data analysis concepts for APC assessment 

are discussed. 

-Finally the conclusions and DLR’s plans for the future are given. 

414 
ntnnn Mr Luft- und Raumfahrt e 3 

o Aircraft-Pilot Coupling (APC) is a highly adverse man-machine problem 

due to disharmonic pilot control inputs. 

o The meaning of the acronym PI0 was changed frompilot-induced 

oscillation to pilot-involved oscillations in order not to blame the pilot. 

o Non-linear effects in the flight control system can cause APC problems 

(f/ying qualities cliff). 

o The APC phenomenon contains three main elements: the pilot, the 

aircraft, and the trigger. 

o APC is no pilot failure, but a failure in the flight control system design 

process. 

. I 

PI0 Workshw, NASA Drvden Fliqht Research Center, Edwards, CA. 6-8 Mril 1999 

- The above list contains the most important key words when talking about 

APC. 

-There is a strong agreement that APC is a highly adverse man-machine 

problem due to disharmonic pilot control inputs. 

-The expression APC was introduced to replace the acronym PI0 first. 

Today APC has a more general meaning than PI0 

-We all know well that nonlinear effects in the FCS can trigger APC. This is 

commonly illuminated by the FQC metaphor 

-Further more we can state that an APC contains 3 elements: pilot, a/c and 

trigger. Pilot is obvious, since without the pilot in the loop no APC is 

possible. The a/c is represented by the complete Flight Control Systems. 

The trigger can have different forms, such as NL-effects, or increased task 

elements, but always causes a sudden change in the closed loop dc-pilot 

system dynamics resulting in a misadaptation of the pilot. 

-Last but not least: APC is no pilot failure, but a failure in the flight control 

system design process. 

415 
cat I PI0 

nonlinearities 

ht Research Center, Edwards. CA, 6-8 Aprl 1999 

This diagram shows a simple classification (not complete). We can see safety 

critical and not safety-critical types of APC. 

Not critical: We have e.g. the low amplitude-high frequency oscillations 

bobbling and ratcheting 

Critical.: Distinguish between non-oscillatory and oscillatory (were we have 

PI0 three categories) 

416 
Dtutsd_s bntrum fOr Luft- und Raumfahrt e.V.

Aircraft-Pilot Coupling (3)

Rate saturation is the dominating nonlinear effect in modern

flight control systems triggering APC (Category II PIO)

Ude I

t

J

time

time delay

PIOWorkshop, NASA Dryden FlightResearchCenter. Edwards. CA. 6-8 April 1999

-The histoD' of aviation has shown that Rate Saturation is the dominating

nonlinear effect in modern flight control systems triggering APC (Categou '

11 PIO).This was the background for defining an individual category for

APC caused by Rate Limiters > category 11PIO.

-The major problem with Rate Saturation is that an additional timedelay is

introduced after Rate Limiters onset. The further point is that this additional

delay is not constant but anaplitude dependent.

417
Deutscha Zentrum for bah:- und RaumMhrt e.V.

_i _

:Prediction of APE (1)

i ! i i_ i

The main objective is to predict potential APC problems in the design

phase of the flight control system.

For that task

several APC prediction criteria are available, such as NeaI-Smith, Bandwidth, Phase

Rate, Smith-Geddes,

and

a comprehensive handling qualities data base is available, such as the flight test

programs NeaI-Smith, LAHOS, HAVE PIO, HAVE CONTROL,

but

most criteria and data bases only address linear effects due to filters and time

delays in the flight control system causing a high frequency phase rolloff.

PIO Workshop, NASA Dryden Flight Research Center. Edwards, CA, 6-8 April 1999

The objective of this presentation is to discuss means and methods used to

predict potential APC problems in the design phase of the flight control

system.

For that task several APC prediction criteria are available, such as NealSmith, Bandwidth, Phase Rate, Smith-Geddes.

But most criteria and data bases only address linear effects due to filters and

time delays in the flight control system causing a high frequency phaserolloff. The high frequency phase-rolloff is the main effect causing category I

PIO.

418
Det_L_hesZentpJmfor Luft- und Ratm_i_rt e.V.

Prediction of APC (2)

Implementation of Rate Limiters in Flight Control Systems

Feedback loop:

- Protectingtheactuatorsagainstoverload

- Defining the maximum rate independent of the flight condition

Forward path;

- Preventing a saturation of the feedback loop limiters due to high pilot

input rates

PIOWorkshop, NASA Dryden Right ResearchCenter, Edwards,CA, 6-8 April 1999

But what about categoD' 11 ?

Let us first have a look at typical implementations of Rate Limiters in

modem FCS. We have two typical locations: In the feed-back loop and in the

foravard path.

In order to predict APC due to these Rate Limiters we have dex, eloped the

OLOP criteria at DLR,

419
I

Deutsch_ ZenUum fGrLufto und Itaurnfshrt e.V.

The OLOP Criterion (1)

OLOP means the Open Loop Onset Point of a rate limiter in an aircraft-pilot loop,

which is plotted in a Nichols chart.

OLOP is a criterion to predicthandling qualities problemsdue to rate limiting in

the flight control system(category II PIO).

OLOP is applicable to the roll, pitch and yaw axesfor rate limiting elements in the

forward path or in the feedback loop of the flight control system.

OLOP hasbeen developed by DLRbased on the describing function technique; the

intensity of the jump resonanceis highly dependent on the OLOP-Iocation.

... TheOLOPcriterion hasall the hallmarksof thepresent author's methodology

for practical design guidance...

John Gibson, 1999

J

PIOWorkshop, NASA Dryden FlightResearchCenter, Edwards,CA, 6-8 April 1999

OLOP means Open Loop Onset Point.

The OLOP criterion is capable to predict category II PIO due to rate

saturation effects.

It is applicable to all rtelated problems.

OLOP has been developed, based on the Nichols amplitude/phase diagrm It

has been shown that the intensity of the jump resonance due to Rate Limiting

onset is highly dependent on the OLOP-Iocation in a Nichols chart. For

OLOP application no Describing Function technique is required.

420
The OLOP Criterion (2)

Validation of the OLOP Criterion

o Flight simulator experiments

on FFA's ground based

simulator FOSIM'.

o Five experienced test pilots

performed 342 simulator

runs.

o DPIOR means the difference

between linear and nonlinear PIO ratings; all runs

were done with and w/o rate

limiting.

o Significant correlation was

found between the DPIORs

and the OLOP criterion.

15

410

E. s

0

-5

-10

-15

_ i _ _ _i_::_i_....

&

• 4 OAO_ • • ,11_

• • o . ._J_O v"

t._ o o

o •

_1 LATHOS F-18 YF-16

DPIOR o 0 z_

DPIOR 2 o 0 A

DPIOR 2 - 3 • • •

DPIOR 3- 4 • • •

-180 -160 -140 -120 -100

phase, deg

"FOSIM: Forskningssimulator

PIO Workshop, NASA Dryden Flight Research Center. Edwards, CA. 6-8 April 1999

Here some high-level information about OLOP are given:

OLOP has been validated by special simulator experiments

FOSIM simulator was used within a collaboration with the Swedish FFA.

342 test runs (using different configurations in the roll axis based on

LATHOS, F-18, YF-16 test pilots) with five test pilots were made.

The results are shown above.

You can see a significant correlation between the OLOP location and the

DPIORs

It is important to correlate the DPIORs with OLOP since OLOP only predicts

APC due to Rate Limiters effects. It is not correlated with the category I P10

criteria.

421
DeutschesZentrum (Grtuft- und I_umfahrt e.V.

The OLOP Criterion (3)

Determination of OLOP

1 Linear frequency response

from stick input to attitude:

•._ pilot model gain

2. Linear frequency response

from stick input to rate limiter

input:

•_ onset frequency u)onset

3. Linear frequency response of

open loop system (loop

opened at the rate limiter):

-)_ OLOP: [phase,gain]@ _Oonset

Closed Loop Aircraft System

rate limiterinput

stick _de

Open Loop Aircraft-Pilot System

PIOWorkshop, NASA Dryden FlightResearchCenter, Edwards,CA, 6-8 April 1999

For OLOP applicaation three linear frequency responses are required.

1. From stick to attitude (this is also required for NeaI-Smith or Bandwidth

criteria) used for the pilot model

2. From stick to rate limiter input > Omega-onset

3. Open loop system including pilot model.

422
Deu_ Zentrum for Luft- und II_umfah_ e.V.

The OLOP Criterion (4)

Influence of Pilot Model Gain

15

• feedback loop rate limiter

•ore.10 " _ ..... " ....... •

forward path • .,j

_. 5 ra I

E

¢¢1 III1_

_11111

0 •

-5

o _= =-160 deg

D _= =-150 deg

-10 o _, = -140 deg

A _= = -130 deg

v ,=,,= -120 deg

-15

-180 -160 -140 -120 -100

phase, deg

PIOWorkshop, NASA Dryden FlightResearchCenter, Edwards.CA, 6-8 April 1999

One special chapter is the pilot model. It is proposed to use simple gain

m(xleis based on Lhe crossover phase angle F,c. Further more a range of pilot

gains should be investigated..

There are two example configurations, one with Rate Limiter in the feedback-loop and one with Rate Limiter in the forward path. This is categor3, II

PIO prone only for very high pilot gains, which means aggressive pilots. The

other configuration (RL in FB-Ioop) is category I1 PIO prone for the entire

pilot model gain range.

Here we will probabl3, have a problem.

42.3
I

Dlml_hes Zentmm for Luft* und Raumfahrt e.V. t=

TheOLOP Criterion (5)

Documentation

o Duda, H.: Effects of Rate Limiting Elements in Fh'ght Control Systems - A New PIO- [

Criterion, AIAA-Paper 95°3204, 1995.

o Duda, H.: Prediction of Pilot-in-the-Loop Oscillations due to rate saturation,

Journal of Guidance, Navigation, and Control, Vol. 20, No. 3, 1997.

o Duda, H.: F_ing Qualities Criteria Considering Rate Limiting, DLR-FB 97-I 5, 1997.

o Duda, H., Duus, G.: New Handling Qualities Database on PIO due to Rate

Saturation, DLR-FB 97-53, 1997.

o Duda, H., Duus, G., Hovmark, G., Forssell, L: New Flight Simulator Experiments

on PIO due to Rate Saturation, AIAA-Paper 98-4336, 1998. [

o Duus, G., Duda, H.: Analysis of the HAVE LIMITS Data Base using the OLOP

Criterion, to be presented at the 1999 AIAA-AFM Conference.

PIO wodr31_o, NASA Dryden Right Research Center. Edwards. CA, 6-8 April 1999

Here a list of the most important documents

- 1995 was the first, where the idea was presented, but the criterion was not

fully developed and no data base was available.

-A very extensive report is this one, but in German

-The next papers describe the data base

-And finally we analysed the HAVE LIMITS data base. The results are

presented at the 1999 AIAA conference in Portland by Gunnar Duus.

424
I11 

Dwtsches &ntrum Mi Luft- und bumfahrt e V. 

Recent Flight Test Experiments with AnAS (1) 

Objectives 

o Final Validation of the OLOP 

criterion using flight test data. 

Identification of pilot model gains in 

the pitch axis. 

oTesting automatic code generation 

tools for software implementation 

on the AllAS experiment computer 

(Sirnulink Real-Time Workshop). 

o Improving flight test evaluation and 

analysis techniques for APC 

assessment. 

b d 

v, PI0 Workshop, NASA Dryden Fliqht Research Center, Edwards. CA. 6-8 April 1999 

The ATTAS experiments: 

There were three objectives: 

Although we consider the OLOP criteria as ready we wanted a final 

validation, especially to get some more experience in the pitch axis. 

We did all the design and analysis work in the Matlab/Simulink environment, 

check Real Time Workshop. Last but not least we plan to develop further our 

flight test data analysis concepts for APC assessment. 

425 
We designed the experiment based on a set of criteria. I will concentrate mq 

talk on the pitch axis. but we did the same thing in the roll axis too. 

In the pitch axis we used the N/S and C* criteria in order to define the linear 

system dqnamics and OLOP for the beha\ iour after Kate Limiters onset. We 

defined baseline confip. one in L1 and one in L2/3. This is depending on the 

band ~idth (BW) \\hen N/S is applied. For this type of a/c BW of 23 is most 

relevant. For in\ estigation of Rate Limiter effects \\e applicd 3 max. rates (7, 

13 and 30 de&) for the elevator deflection. 

The diagram shoirs see the OLOP locations. It is interesting. that nith 

increasing ma\. rate the category 11 PI0 potential seems to he bigger. This is 

a point where ne were not able to clarify this bj the flizht test results. We 

assumed a time delq responsible for this result. 

426 
lIS 

D.utrchm Lentrum fllr Luft- d Rannfahrt e.V. 

xperirnents with AlTAS (3) 

oftware Implementation via Sirnulink Real-Time Workshop 

v, PI0 WorkshoD. NASA Dwden Aiqht Research Center, Edwards, CA, 6-8 npril 1999 

This diagram depicts our sh implementation concept. We del eloped simple 

controllers under Simulinh. In the pitch avis it is nz or C* law, containing q 

and nz feedback and one integrator. 

llsing the Real Time Workshop we simpl) pushed a button and got ;1 C-code 

\\ hich is implemented on the ATTAS euperiment computer. 

This is a verj exciting technique nhich ne did first time for these 

experiments. Quite a lot of sh adaptation norh was required. hut ne non 

ha\e a exellent basis for future euperiments. 

427 
lb 

Recent Flight Test Experiments 

o Software implementation via /?ea/- 

Time Workshop works well and 

' 

provides a very good basis for future 

experiments. 

osignificant correlation between pilot 

comments and predictions based on 

the criteria was obtained. 

Category II PI0 in the pitch axis for a 

basically stable aircraft. In the roll 

axis Category II PI0 is more likely. 

oPilot gains were much smaller than 

expected, especially in the pitch axis. 

olt is very "difficult" to produce a 

,s-c-,*Wro". *w)* 1 1 I I I I 

dq A 

I- - - 

-7 s*k&& ' ' ' ' ' ' ' 

Is 

-; ' ' ' 

Dm. K 

This chart shows the main experiment results: 

First the sh implementation was greatly facilitated using Real Time 

Workshop. 

A significant correlation between pilot comments and predictions based on 

the criteria \viis obtained 

A very interesting result is, that it is .'difficult" or very unlihcl!, to get 

category I1 PI0 in the pitch a\is with stable aircraft. 

There is one esample - a run with a max. rate of 7 deg/s. which is very low. - 

The pilot gave a PlOR of 1-2. Here is one esplanation: The depicted example 

shmvs a tracking task with a commanded pitch angle. Pilot activities show 

that the pilot gains tverc much smaller than expccted. I \vi11 come back to this 

point later. 

428 
II

Deutsches Zentrum fGr Luft- und Raumfahrt e,V.

OLOP Evaluation of two HAVE LIMITS Configurations

2O

rate J_mitldeg/sec] ' ' \'

_3o '_lo

•_ _50 , .-"

ElC_" _60 "'_o20

ro _ 4 i

.d- o3C \ 2DU ,d_ -c_°40

5 --_ 50

IIIIIIII1_ / _"

0

57 # HAVELIMITS

-s ,"" _ 2DU: unstable a/c

,_ ,_.2D: stable a/c

"'#'(_)0 -180 -160 -140 -120 -I00 -80

phase, deg

PIO Workshop, NASA Dryden Right Research Center, Edwards, CA, 6-8 April 1999

J

_60

Here is one more chart to confirm the statement that category ll PIO for

stable a/c is very unlikely - the HAVE LIMITS program (to be presented on

AIAA 1999).

You see two configs, from HL evaluated with the OLOP: 2D represents a

stable a/c, while 2DU represents an unstable ale. 2D runs into the dangerous

area only very low Rate Limitations, while 2DU is category II PIO prone

even for quite high max. rates.

]"his result is well in-line with the PT results obtained in the HAVE LIMITS

program. Gunnar Duus will give more details on this study in Portland.

429
Deutsches Zentrurn for L.uft- und Raumfahrt e.V.

Data Analysis Techniques for APC Assessment (1)

The objective is to develop procedures for APC-Assessment based on flight test

data complementary to the pilot ratings.

Approach

o Identification of simple aircraft and flight control system (FCS) models from the

flight test data.

o Evaluation of handing qualities criteria using the identified aircraft and FCS

models.

o Comparison of criteria results with pilot comments.

o Identification of pilot models for the evaluation of the OLOP criterion.

PIO Workshop, NASA Dryden Flight Research _:enter, Edwards, CA, 6-8 Al_ii 1999

Now I come to the data analysis. The objective is to develop procedures for

APC-Assessment based on flight test data complementary to the pilot ratings.

The pilot rating is always subjective and it is quite easy not to find a "hidden

weakness". So numerical data analysis is an important factor in order to

maximise flight safety.

Our approach is to identify simple a/c- and FCS- models and evaluate

Handling Qualities criteria and compare the numeric results with the pilot

comments.

Furthermore we identify simple pilot models for application of OLOP.

430
DeutschesZenUum fGr Luft- und Raumfahrt e.V.

Data Analysis Techniques for AP( Assessment (2)

identification Concept

a) Fourier transforms from stick

input to aircraft output signals;

approximation of transfer functions

b] Linear aircraft models in the time

domain from control surface

deflection to aircraft output signals.

c) Linear aircraft-FCS models in the

time domain from control stick input

to aircraft output signals.

d) FCS time delays using the results

from b) and the known FCS gains

and rate limits; to be used for OLOP

evaluation.

surface

stick input deflection aircraft output

flight control system aircraft

FCSfeedback signals

PIOWorkshop, NASA DrydenPlightResearchCenter, Edwards,CA. 6-8 April 1999

I will now discuss different concepts for a/c-FCS mode identification.

The first one works in frequency domain. Transfer functions are

approximated to the fast fourier transforms of the test data.

Method b) is only required for d): it means the identification of linear a/c

models using surface deflection as input and a/c reaction as output.

Method c) uses stick signals as input. An equivalent time delay is estimated.

For method d) only delays in the forward path and feedback loop of the FCS

ore identified, while the FCS gains, the maximum rate of the limiters and the

linear a/c models are fixed.

This technique is required to evaluate OLOP from FT data. OLOP can not be

evaluated correctly based on method a) and c) (exception: rate limiters in the

forward path).

431
Deutsdms Zentrum fi_ Luft- und Raumfahrt e.V.

Data Analysis Techniques for APC Assessment (3).

Identification of Aircraft-FCS Models, methods a) and c)

a) FrequencyDomain Identlfl(atlon

dmm_,lu_e _pit_ r_ per s_ck det/ectmn}

_o_

18

-1SO.phase (pitch rate Imr s_k _ll_t,i_)

7t

_Joo

-_soo, I

.:..--L " ........... m_

0

• J ,.0._

o_z o.4 o.5 u 1 2 • s g io

m. riKI/s f...fast fourier uansforms 1

l.-- aoo.ox_a_on j

:/ii _ !i _

0 Time Domain IdentlficatJon (Equivalent Model)

0s p_;_ st_ der_b_

i 2 I 4 s i 7 8 9 io

/........ 1

PIOWorkshop. NASA D_den FlightResearchCenter. Edwards. CA, 6-8 A,or__999

On this chart methods a) and c) are illustrated.

Right: Method a) is a little bit more difficult to apply, you have to decide

about the frequency range to be considered. In this case we did the

approximation up to a frequency of eight rad/s.

Left: Here you see the identification of an equivalent linear model. Here we

have a 3211 input signal, so that it is difficult to include the phugoid motion

due to the short time of the run.

It has been shown that an PID of the tracking task (duration = 120 s) is

favourable.

432
Deutsches Zentrum fGr Luft- undRmamfahrt e.V.

Data Analysis Techniques for APC Assessment (4)

Identification of FCSTime Delays, Method d)

pitch stick deflection

0.5,' _ 1 ,

0 _ .............

n_:l i I L I t

"v"5_ pitch ",,ate , ., , , ..

deg O_

_5t i I i r " t

0 1 2 3 4 5 6

f-measurement 1 time, sec

model w/o FCS delay

model with FCS delay

PIO workshop, NASA Dryden Flight Research Center. Edwards, CA, 6-8 April 1999

This chart shows one PID result of concept d)

The red curve represents the a/c-FCS model response without time delay.

The blue curve the response with time delays.

You see that we have a better matching with delay.

433
Deutsd_s Zantmm f_r Luft- uml Raumfahrt e.Y.

Data Analysis Techniques for APC Assessment (5)

Comparison of Different Identification Concepts

NeaI-Smith Criterion, (o=.= 2.5 rad/s

M

-o Level3 [

10 o conf. I (PIOR 1-21

o o conf. 2 (PIOR 4)

o_-6_ _ Predictions

"_ 4_ , Q_ I• • Identification in frequency

"_ 0__. Level 1 • • Identification in time domain, equivalent models, method c)

=

-2 '- i • Identification in time domain,

, = only time delays, method d)j

-40 -20 0 20 40 60 80

pilot compensation, deg

PIO Workshop, NASA Dryden Right Research Center, Edwards, CA, 6-8 April 1999

This chart shows the results of the three Identification concepts for the pitch

axis configs. Additionally we see the predictions based on the model and

assumed time delay we used before FI'. The main cause for the difference

between Identification and prediction is the assumed delay.

For config 1 we got very consistent results, but we have some scattering for

config 2. This is because this configuration is quite sensitive to additional

delays,

Method d) (only ientification of delays) provides the most consistent results

compared to the pilot ratings. However we are not quite clear alyaut this

config. We need to do some further analysis and PT.

434
DemtschuZ_ntmrnforLuft-emdlitmumfahft e.V.

Data Analysis Techniques for APE Assessment (6)

Identification of Pilot Model Gains

:3

Approach

Parallel simulation of the

closed loop aircraft pilot

system.

Manual adjustment of pilot

gain in order to get

"similar" closed loop

performance, such as

damping and overshoot.

Results

Crossover phase angles for

all configurations:

pitch axis:-90 to -100 deg

roll axis: -110 to -120 deg

PiOWorkshop,NASADrydenFlightResearchCenter,Edwards,CA.6-8April1999

For the evaluation of OLOP we need simple pilot models. For that purpose

we do a parallel simulation of the closed loop a/c-pilot model. The input

model gain is adjusted manually in order to get "similar" closed loop

performance, such as damping and overshoot.

In this case we got crossover phase angles significantly lower than expected.

For experiment design we assumed -130 deg as medium gain.

In the roll axis this is slightly higher.

435
Zentmm fGr Luft. und bumfahrt e.V.

Data Analysis Techniques for APC Assessment (7)

OLOP Criterion, Conf. 1, _, = -100 dig

15t 1 .,,_

"0 _ I •

1 -

°:!I

;3o

f

-180 -160 -140 -120 -100 -80 -60

phase, de9

• PIO Workshop, NASA Dryden Flight Research Center. Edwards, CA, 6-8 April 1999

The identified a/c-FCS and pilot models are used lk_r evaluation of the OLOP

criterion. This chart shows config 1- the predicted and identified model for

different max. rates.

You see that OLOP does not predict any categoD, 11PIO problems, which is

well in-line with the pilot comments. The pilot rated this config with P1OR 1-

2 for 30 and 7 deg/s max. rate.

We did not fly the 13 dens case.

436
DlmtKhu hntmm mr Luft- und Rmmlfahn e.V.

Conclusions

o Flight test experiments with AI-IAS were conducted in order to

improve the knowledge base on the OLOP criterion, to test new

software implementation procedures and flight test data analysis

techniques.

o The pilot comments obtained are correlated with the predictions

of the criteria (OLOP, NeaI-Smith).

o Software implementation via Real-Time Workshop (Simulink)

works well and provides a good basis for future experiments.

o Different concepts for flight test data analysis were evaluated; the

OLOP criterion was successfully evaluated on the basis of the

identified aircraft and flight control system models.

PIO Workshop, NASA Dn_len Flight Research Center, Edwards, CA, 6-8 A,orit 1999

i

Conclusions:

We did Flight test experiments with ATTAS in order to improve the

knowledge base on the OLOP criterion especially in the pitch axis, to test

new software implementation procedures and to improve flight test data

analysis techniques.

The pilot comments obtained are correlated with the predictions of the

criteria (OLOP, NeaI-Smith).

Software implementation via Real-Time Workshop (the C-code generator of

Simulink) works well and provides a good basis for future experiments.

Different concepts for flight test data analysis were evaluated; the identified

aircraftand pilot flight control system models.

437
I

_tsdlls _t.qJm for IJJ_- utKl b_|lh_ e.V.

Future Activities

I I I

The flight test experiments presented have prototype character; the work

is going to be continued with respect to

o Experiments with more APC prone configurations, such as aircraft

with relaxed static stability.

o Testing of on-line APC detection and warning algorithms.

o Evaluation of phase compensation filters in order to reduce the time

delay due to rate limiting.

o APC demonstration maneuvers.

Long Term Objective

A standard for APC testing o f highly augmen ted aircraft

,,, , ,J

PIO workshop, NASA Dryden Flight Research Center. Edwards. CA. 6-8 April 1999

438
Criteria to Simulation to Flight Test

- and Vice Versa

David G. Mitchell

Technical Director

Hoh Aeronautics, Inc.

Pilot Induced Oscillation Research Workshop

NASA Dryden Flight Research Center

7 April 1999

Outline

• Steps for minimizing PIO risk

• Assessing risk if a PIO occurs

• A possible PIO rating system

• Pilot variability in PIO simulation

• Some recommendations

439
Steps for Minimizing PIO Risk

1. Be prepared for PIO

2. Apply criteria to design

3. Use criteria to focus preliminary simulations

4. Use early flight data to update sim. model

5. Repeat steps 1 - 4

6. Use simulation to apply criteria for large inputs

7. Use criteria to focus preliminary flight tests

8. Use real-time onboard detection for early warning

9. Repeat steps 1 - 8

Be Prepared for PIO

• Military procurements represent a dichotomy:

- Projects adopt success-oriented scheduling

- Evaluators expect to encounter PIO in flight test

• PIOs will almost always occur

- Should not be a surprise

- Testing must be adopted to look for them

• The more advanced the aircraft (unstable, multiple

effectors, multi-purpose effectors, complex

augmentation) the greater the potential for

catastrophic PIO

440
Be Prepared for PIO (concluded)

• Pilots must be a part of the process

- Familiar with the phenomenon

- Aware of potential through all phases of testing

• PIO is not an operationally relevant event

- Test pilots' job is to go beyond normal operations

- If test pilot won't push the airplane, rest assured that some

unsuspecting tleet pilot will

- Any flight test can be a test for PIO tendency

• If a PIO occurs, there must be a way to assess risk of

continuing flight testing before a fix is found

Steps for Minimizing PIO Risk

1. Be prepared for PIO

2. Apply criteria to design

- As early as possible in design process

- If you apply valid criteria and your airplane fails, it

doesn't mean the criteria are bad

3. Use criteria to focus preliminary simulations

4. Use early flight data to update sim. model

5. Repeat steps 1 - 4

6. Use simulation to apply criteria for large inputs

7. Use criteria to focus preliminary flight tests

8. Use real-time onboard detection for early warning

9. Repeat steps 1 - 8

fli L-""

441
Steps for Minimizing PIO Risk

1. Be prepared for PIO

2. Apply criteria to design

3. Use criteria to focus preliminary simulations

- Don't spend time in areas where criteria are easily met

- If criteria predict PIO -- fix the design!

4. Use early tlight data to update sim. model

5, Repeat steps 1 - 4

6. Use simula_on to apply cnteria for large inputs

7. Use criteria to focus preliminary flight tests

8. Use real-time onboard detection for early warning

9. Repeat steps 1 - 8

Steps for Minimizing PIO Risk

1. Be prepared for PIO

2. Apply criteria to design

3. Use criteria to focus preliminary simulations

4. Use early flight data to update sire. model

- It should contain all known nonlinearities and limits

5. Repeat steps 1 - 4

6. Use simulation to apply criteria for large inputs

7. Use criteria to focus preliminary flight tests

8. Use real-time onboard detection for early warning

9. Repeat steps 1 - 8

442
Steps for Minimizing PIO Risk

1. Be prepared for PIO

2. Apply criteria to design

3. Use criteria to focus preliminary simulations

4. Use early flight data to update design model

5. Repeat steps 1 - 4

6. Use simulation to apply criteria for large inputs

- Frequency sweeps to control limits

- Even if sim. is doubtful for PIO, it can be useful for

applying inputs beyond those considered safe in flight

7. Use criteria to focus preliminary flight tests

8. Use real-time onboard detection for early warning

9. Repeat steps 1 - 8

Steps for Minimizing PIO Risk

1. Be prepared for PIO

2. Apply criteria to design

3. Use cdteria to focus preliminary simulations

4. Use early flight data to update design model

5. Repeat steps 1 - 4

6. Use simulation to apply criteria for large inputs

7. Use criteria to focus preliminary flight tests

8. Use real-time onboard detection for early warning

9. Repeat steps 1 - 8

443
Steps for Minimizing PIO Risk

1. Be prepared for PIO

2. Apply criteria to design

3. Use criteria to focus preliminary simulations

4. Use early flight data to update design model

5. Repeat Steps 1 - 4

6. Use simulation to apply criteria for large inputs

7. Use criteria to focus preliminary flight tests

8. Use real-time onboard detection for early

warning

- Tomorrow morning

9. Repeat steps 1 - 8

Assessing Risk if a PIO Occurs

• If PIO occurs in the development process, it must

always be treated with concern

- Fix the problem!

• It may be necessary, and possible, to continue the

development effort

• Risk is a function of several factors:

- Category of PIO

- Severity of PIO

- Frequency of occurrence and duration of PIO

444
Reducing Risk: Categorize the PIO

• Category I (linear):

- it should be possible to quickly identify causal factors

- Lowest risk to continued operation

• Category II (rate limiting or other saturation):

- More difficult to identify causes

- Risk depends on other factors:

• Flight condition/aircraft configura_on -- avoidable?

• Consequence of saturation -- unstable airplane?

• Category III (nonlinear with mode switching):

- Highest risk, factors similar to Category II

Current PIO Tendency Rating Scale

• Problems with scale

- Does not mention "tendency"

- PIOR = 2, 3: not relevant to PIO

- PIOR = 4: no indication of severity

- Attempts to mix handling qualities

with PIO assessment

• Examples:

- Pitch bobble (PIOR = 4) with

inadequate control power (HQR = 8)

- Severe (but not divergent) PIO

(PIOR = 4) that is unacceptable

(HQR = 8)

1

No

Yes

Yes

t -...° i

Abrupt Maneuvers

or

_ght Co_rol

_ Yes 6

I pilol Affempls I

to Enter Control

Loop

445
A Possible PIO Rating System

Severity Frequency of Demands Overall

occurrence on pilot assessment

Dangerous Never stopped Couldn't prevent it What airplane?

(bail out) (abandon airplane)

Severe Couldn't prevent it Intolerable for

(abandon task) Most of the time (Abandon task) the task

(fix it)

Prevented or alleviated

Moderate by technique Objectionabte

(can't ignore it) Occasional (task performance (warrants

compromised) improvement)

Prevented or eliminated Tolerable

Mild Only a very short time by technique (satisfactory without

(can ignore it) (task performance improvement)

not compromised)

None Never saw one No tendency to What PIO?

_1 j_,lb.==_= induce oscillations

PIO Rating System Allows for Risk

Assessment in the Development Process

- Example: PIO Severity vs. Frequency of Occurrence

Frequency of occurrence

Never Most of the Occasional Only a very Never saw a

time short time PIO

Dangerous

(bail out)

Severity Severe

(abandon task)

Moderate

(can't ignore it)

Mild

(can ignore it)

None

High High High

High Moderate

Moderate Moderate Low

Low Low

446
Pilot Variability

• Variability in pilot opinion is well-documented in

handling qualities experiments

- Test pilots have varying backgrounds, expectations, flying

styles

- This is good! Fleet pilots will be even more diverse

• Variability is magnified when it comes to PIO tests

and exposure of PIO tendencies

• Monitor pilot performance for tracking tasks

- Expect variability in performance (example: recent sim.)

Pilot Variability in PIO Simulation

• Example: HAVE LIMITS Config. 2DU, 20-deg/sec RL, discrete

tracking task, flown on USAF LAMARS simulator

• Some (minor) differences in setup between sim. and flight

• Results below are typical of sim. (10 pilots total)

- Different pilots encountered PIO at different rate limits

Facility

NT-33A (Flight)

LAMARS

(Moving-base

simulation)

15ilotI.D. HQR PIOR

1 10 6

2 10 6

10

A 10 5

B 10 5

C 10 6

E 10 6

F 10 5

H 10 5

447
Pilot Variability in PIO Simulation

Plot shows measured crossover frequency (q/qerror) from

discrete tracking task vs. total run time

- Task started at t = 10 sec, ended at t = 138 sec

- Run ended if pilot encountered rapidly divergent PIO

2.5

1.5

Appto_

aomo_r

Ir_uenof

(racZ._ec)

11

_tatA •

C

F

Eo H

B

RlatD

(Compl_d Task)

20 40 60 80 100 120 140

I_zmIwn_tl (me)

Pilot Variability in PIO Simulation

Ten-second sample of long. stick for two highest-crossover

pilots (A and C) and two lowest-crossover pilots (B and D)

- Pilots A and C consistently show larger, more rapid inputs

2-5

2 /\

•- / ' _._ ' iJ!_ ' 'l

O.s ',.- ! /-\ _ ; 3 _\ , '

:; lime I

V

0

..O5

I,II _11

448
Amplitude of PIO

Monitor time-history data for evidence of PIO

- Pilots aren't always aware of PIO on simulator

- Events that seem mild to the pilot may be severe in flight

- Work with the pilot as much as possible!

HAVE PIO Rating Comparisons: PIOR

5

"4

E_

t-

_3

2

1

jff _"

//jl "f_

o ,-)- • y'-_ " _ ,.""

AIO/ 4- "_7 /

• ,_o /// / ///

AO _ // / ///

// / //

/ /

/ / '/,dr0 0 • /////

/J /./"

// "4) / /////I ....

",0, J .._,i" I' VIVIS_rge r_ion

2 3 4 5 6

Mean PIOR Simulation

449
HAVE PIO Rating Comparisons" HQR

10 = -" I A • /. /

/ j

9 -.t-,,,I i// / x,/ ..

8 + ie.." / i.

• + .z / /i

u. -t-_ve/ i

6

"I5 /4/

i I q_Ill(" t /

4 4- @" .,ll/Ail A"

3 ,, I_

/ .,_ .'"'% • I" VMS Fixed base

2 //_ -_- -_, l+ MS_lFixed base

11 I I I i I I I

1 2 3 4 5 6 7 8 9 10

Mean HQR Simulation

1

PIO May Be More Severe in

Simulators

20 -

A

15 -

Blacklines: flight ,, _,

program (Pilot A, _o- ,, , ' IPIOR6, HQR10) ., ':":: ,; ii '<-i_.,?,. i ;.

pitch 5 - . _ _; ,, _ , _s_ _.i' r ,._

Red and blue rate _ i,j_.._,_.fA _i,_llt ,_l.-

lines: MS-1 (d_,,e=) 0 o_,. _ ,_. __:_,Nsimulation (Pilot 2, _,' ,'+ _', _/.- '" " ,, _..... "

two sessions, -s- " '_."_ • "_] "i"_ " V)|

PIOR 4, HQR 6) 40 - ,_

7

40

time (sec)

-15 -

450
Recommendations

• Make maximum use of criteria, simulation, and flight test

• Simulation has value as an adjunct to flight

• Be prepared for PIO

• Assess risk for continuing if PIO is encountered in the

development process

• Expect pilot variability

• Look at both qualitative and quantitative information from

simulation

- Ratings tend to be better

- PlOs may be more severe

451
q
Appendix 3

453
q
Designing to prevent safety-related PIO

PIO Workshop, NASA Dryden, 6th - 8th April 1999

J C Gibson

British Aerospace Warton (retired), Consultant

Introduction

Though PIO is not a new phenomenon, its current notoriety has been acquired in the past two

decades mainly from the all-too frequent serious and sometimes catastrophic examples exhibited

in fly by wire aircraft. Such severe examples were a rarity in the earlier "classical" aircraft with

conventional control systems. Yet the fly by wire technology had brought with it the power to

provide almost any desired handling response qualities. PIOs and sometimes other handling

problems of the "high order" type (to distinguish them from the usually much less severe "low

order" types possible with conventional dynamics) were actually not generic to the technology as

was commonly believed at one time but were inadvertent artefacts of the control system designers.

Since the PIO characteristics were "designed in", they can also be "designed out".

The intellectual rigour necessary to prevent PIO by design must be spread out far beyond the

discipline of the control law specialists. Section 9 of Reference 1 discusses the team approach

essential for the design and evaluation process, and notes the many failures that have resulted

from neglecting this. The repeated examples indicate that newcomers to the fly by wire field have

found it difficult to believe that the problem could happen to them, and so have not implemented a

meticulous anti-PIO design policy. Safety-related, high-order type PIO is not a problem with no

practical solution, preventable only by good luck. The author's 1978 paper on the Tornado PIO in

1976 and its solution (Reference 2) was greeted with surprise, since it was not normal in the

conference circuits to admit to such a problem even though it was widespread. The latter head-inthe-sand attitude probably contributed to the continuing occurrence of safety-related PIO, and only

more recently was the author's example followed by what is now a flood of data and information

on the problem.

The author's own brush with PIO and its solution led to a design methodology to eliminate it in

future projects. The success of this was demonstrated from the early 1980s onwards by a series

of highly unstable aircraft with digital FBW control, namely the Jaguar FBW demonstrator, the

EAP demonstrator and the Eurofighter 2000. Each took to the air with a growing certainty that

safety-related PIO would not be experienced or even be possible, a certainty that proved to be

justified. The rather simple physical principles of control system design for PIO prevention are

discussed in Reference 3.

Use and misuse of specifications

Designers are very likely to get into trouble if they simply design to satisfy customer

specifications. It is not practical to impose specification criteria for handling qualities design in

sufficient detail to ensure good handling qualities while not unnecessarily restricting other design

possibilities that may actually improve on the classical response types. It is not the business of a

government department to design control systems. Practical specifications provide some "must

have" requirements, but one that tries to cover too much ground at once with too few parameters

risks allowing unsatisfactory behaviour to slip through if it is used as the only design guidance.

Perhaps the best known example is the specification for short period frequency versus n/a. Level

1 handling has never been achieved with frequencies near the upper limit, except for good landing

approach control. The latter is most unlikely with minimum allowable frequencies, but good

handling has been achieved at higher speeds with lower frequencies.

455
Another examplein Figure 1 is from genericASTOVL handlingresearchfor the jet-borne

hoveringphaseon ahigh fidelity motionplatform. Two of thecasesareplotted on an attitude

responsemode criterion from the rotary wing aircraft specificationADS-33C. This criterion

quantifiesthe handlingby the bandwidthandhigh ordereffectsby thephasedelay.Both cases,

assessedin the task of lateraltranslationalcontrol, are nominally secondorder roll attitude

responseswith a bandwidthof 6 radiansper second.Their actualbandwidthdecreaseswith

increasingphasedelay,which wascreatedby an additionalsecondorder lag to representhigh

ordereffects.This genericfourthorder modelformat was derivedfrom a designstudy for the

VAAC Harrierresearchaircraftandrepresentedits highordersystemdynamicsveryaccurately.

However,theresultswerenot whatthecriterionwould leadoneto expect.In casel(a), as the

bandwidthdecreasedwith increasingphasedelay,thetranslationtaskhandlingqualitiesremained

constant.Thesequalitieswere found to be relatedto specifictime responsecharacteristicsthat

remainedeffectively unchangedfrom the baselinebandwidthcase.Therewas an increasing

untidinessin attitudecontrolinducedby the high orderlag, thoughthe effectswere acceptable

overtherangetested.Case1(b) with higherbandwidth,despiteremainingcompletelywithin the

criterionLevel 1 region, deterioratedinto severeattitudecontrol PIO, exacerbatedby lateral

accelerationforceson thestick andpilot's arm with thecockpitmountedon top of the platform.

Thecauselay in thehighPIOgainof theattitudefrequencyresponse,which is not accountedfor

by this criterion. The only differencebetweenthe caseswas that l(a) had a nominal mode

dampingof 1.0and l(b) hadadampingof 0"5.

The criterion broadlyquantifiedthe handlingof Case l(a), but it was misleadingeither as a

contractspecificationor as a designcriterion when appliedto circumstancespresumablynot

envisagedin its original derivation.It is not known if it was testedfor responseswith low

damping,for example,eventhoughthisis permittedelsewherein thespecification.

Potentialdifficulties canbecausedby anyotherlimited-parametercriterion.Figure2 shows the

pitchattitudeNicholsplotsfor theYF-17astestedby Calspan,in theoriginal severelyPIO-prone

form andthe very satisfactorymodifiedversion.To theinformedeye,the badandgood natures

of the respectiveresponsesare instantly obvious from the presenteddetail alone, but it is

necessaryto havesomeformalisedcriteria to quantify this. The modified casewas one of the

smallnumberof exampleswith excellenthandlingaroundwhich the author developedthe socalled"Gibsoncriteria"boundariesin Reference4 from 1982,theonefor landingapproachbeing

shownin thefigure. The boundariesdid indeedcapturemuchof theessenceof good handling,

but werenarrowlyconstrainedandwerelaterfoundto excludeotherperfectlyacceptableresponse

shapes.Similarproblemsarosewith theso-called"Gibsoncriteria"timeresponseobservationsin

Reference4, which againwerederivedfrom a fairly limitedsetof cases.The authoralso learned

thehardway thatsometimesothersof adogmaticframeof mind couldfind it difficult to accepta

responsethatdid notentirelysatisfythe boundaries"becauseit violatesthecriterion", despitehis

protestationsthattheywereintendedasindicativeguidelinesandnotabsolutego/no-golimits.

Neverthelessthesecriteriaappearfrom the literatureto havebeenof assistanceto a numberof

other designers,and were an essentialgrounding to the author's later design methodology

describedin Reference3. In this, there is a much reducedemphasison attitudefrequency

response"shape"boundariesbecausetheyinherentlychangetheircharacteristicswith increasesin

truespeedandaltitude.Thenatureof pitchbehaviourin the"generalhandling"regionof Figure2

is richly illustratedfor designpurposesby time responsessuchasflight pathtime delay,attitude

dropbackand pitch rate overshoot,which cannot be quantifieddirectly from the frequency

responseeventhoughthey maybeobviously presentby visual inspection.On the otherhand,

whilehighorderPIOtendenciesareeasilyobservedby a lagin thetimedomainpitchacceleration

456
response,they are more clearly delineatedin a detailedanalysisof the frequency response

characteristicsin the "safety-relatedPIO" region of Figure 2, independentlyof the general

handling. All this is discussedin Reference3. (Time responsesare an excellentdesigntool,

irrespectiveof theirunsuitabilityfor flight testanalysis.)

A varietyof delaycriteriahavebeenpromoted,of whichphasedelay(or theaveragephaseratein

theauthor's terminology)is themostaccuratemeasureof the actualdynamicsthat may leadto

PIO, particularlyof Type 1 thoughobviouslythesemayin turn leadon into Type 2 or Type 3

PIO. It is doubtfulif suchcriteriahaveany meaningfor analysisof largeamplituderesponses

with non-linearactuationeffects, however. The authorfound it unprofitable to attemptthe

laborioustime responseanalysisfor phasedelayin thisregime.

The primary importanceof phasedelayis to indicatea significant lag in the initial rotational

accelerationtime responseto a pilot's control input which may lead to a Type 1 PIO. If this

divergesinto theactuatorsaturationregime,the PIO continuesat a decreasingfrequencywhich

remainsuniquelyrelatedto the 180degreelag in attitudeasthenon-lineareffectsbecomemore

pronouncedwith increasingamplitude.If on the otherhanda largesaturatedPIO burstsinto life

with nointerveninggrowthfrom smallbeginnings,thenit instantlylocksonto thePIO frequency

in thesameway.In neithercaseis thereanysignificancein therateof phaseanglevariationovera

rangeof frequencybeyondthePIO, which in effect is phasedelay. What does matteris the

mannerin whichtheattituderesponseattheuniquePIOfrequenciesvariesfrom the linearcaseas

thepilot's input amplitudeincreases.

Thehandlingqualitiesspecificationsknown to the authordo not addressthe safety-relatedPIO

problemdirectly,otherthanto requirethat it mustnot occur. Thesespecificationsare generally

assumedto applyto thelinearregime,presumablybecausetheyaremostlyexpressedin termsof

parameterssuited to straightforward frequency response analysis techniques. The few

requirementsspecificallyassociatedwith full amplitudecontrol inputs, which would certainly

invoke any actuationand aerodynamicnon-linearities,are typically open loop time response

requirementssuchasroll performance,and would not necessarilyillustrateany PIO tendency.

Neverthelessthereis no generalexclusionof large amplitudeand non-linear conditions from

consideration,and indeed"the effects of the control equipment should not be overlooked" in

calculations or analyses directed towards investigation of compliance with the specifications.

The realm of the safety-related high order PIO

The following is a brief resume of the author's successful experience in high order PIO solution

and subsequent elimination by design over the period from 1976 up to the present, extracted

mostly from Reference 3.

At the time of the 1976 Tornado landing PIO, there were no criteria or appropriate data generally

available to explain it. However, it had clearly grown out of the stick pumping in the landing

flare, an activity described by Bihrle in 1966. He noted that just before touchdown, pilots would

often engage in a rapid pitch control oscillation in phase with pitch acceleration, at frequencies

well above the short period. The acceleration amplitude was consistently around +_6.5 deg/sec_.

Bihrle concluded that pilots acted this way to generate confidence in pitch control as the speed

reduced towards the stall when very precise flight path control was needed for a smooth and safe

landing. The activity was also quite subconscious, all pilots being unaware of it.

The author had used the stick pumping theory in the Tornado design process to ensure that there

was adequate hydraulic pump flow capacity at idle engine rpm in the landing approach, and in fact

found in flight records that pilots did stick pump as predicted. However, the Tornado pitch

attitude dynamics differed significantly from previous conventional aircraft. These consistently

feature stick pumping at typically 8 to 10 rad/sec resulting in an attitude oscillation that is very

457
small.Theamplitudeis usuallylessthanafifth of adegreepeakto peakandis effectively

unnoticeable.TheTornadostickpumpingfrequencywasabout3 to4 rad/sec,andatthenominal

accelerationleveltheattitudewouldbearound2degreespeakto peak.Somepilots usedlarger

pumpingamplitudesthanothers.Thelikely triggerseemedto bethatthepilot suddenlybecame

awareof theattitudeoscillation,andwaspresentedunexpectedlywithaready-madePIO situation

withtheattitudealready180degreesoutof phase.

Stickpumpingdoesnot triggerPIOin conventionalaircraft.Theobvioussolutionatthetimewas

to ensurethattheattitudedynamicsin thestickpumpingfrequencyregionweremadeto favour

thesubconsciouspitchaccelerationpumpingactivity, andnotto encouragethepossibilityof the

unstablepilot-attitudePIOcouplingwhichoccursat similarfrequencies.The"synchronouspilot"

PIOmodelproposedbyAshkenasandMcRueraround1964,expressedasagainelementand

assumedto applycontrolin anti-phaseto theattitudeoscillation,wasclearlyevidentin the

TornadoPIO.With nopilot phasecontribution,theclosedloopinstabilitynaturallyoccurredat

thefrequencywheretheaircraftattitudephaselagto controlinputswasaround180degrees.The

authorconcentratedstudiesontheaircraftdynamicsin thisregion.

Figure3 showsthecalculatedTornadolandingcasepitchattitudefrequencyresponsesfor four

differentpitchcontrollaw configurations.The unaugmentedmodewasrathersluggishbutwas

otherwiseperfectlyacceptable.It hadalreadybecomeclearthatthestickcommandgainatlow

speedsin thefirst augmentedversion,whichexperiencedthePIO,wastoo highasit was

excessivelyeasyto saturatethepitchcontrolsystem.Thelargeamplituderatioatthe 180° phase

lagfrequencymeantthatlargeoscillationscouldeasilybegeneratedby quitemoderatestick

inputs.In thecompleteabsenceof anyothercriterionwhatever,thepolicy wasadoptedthata

stabilitymarginmustremainif anypilot againusedthesamegainasin theaccident.

Thesecondcontrollaw version,whichwasnearlyin aflight clearedstatusatthetimeof the

accident,hadalreadyhalvedthePIOresponsegainatlow speedswith its substantialreductionin

stickcommandgain,andwasapprovedfor use.Theauthorexpressedreservationsbecausethe

lineardynamiccharacteristicsof thesecondversionwerelittle changedfrom thefirst version.The

sensationpilotshadof havingto "feelfor theground"in thefirst versionwascausedby amarked

lag in theonsetof pitchaccelerationin thetimeresponse,whichwasmuchlargerthanin the

unaugmentedcasewhereconventionalactuatordynamicsweretheonly highordereffect.In the

secondversionthetransientaccelerationlaghadbeenscarcelyreducedat all,andsomepilots still

founda slightimprecisionattouchdown.Theauthor'sconcernwaseventuallyjustified by an

incipientnon-divergentPIO,distinguishedin theflight recordmainlyby thepilot's statementthat

hehadsensedits onset.As thetailplaneswerecloseto theirnominalratelimit, theeffectivesafety

marginwasunacceptablysmall.Furtheruseof full augmentationfortakeoff andlandingwas

againprohibiteduntil afinal solutionwasdeveloped.

The third versionfollowed the author's embryonicideasabout the importanceof the attitude

dynamicsaroundthe 180degreephaselag frequency.It further reducedthe PIO gain and the

transientaccelerationlag by speed-dependent schedulingof thelag-leadstick commandpre-filter

to a unitygainat low speed.Thelag-leadwas restoredat higherspeedsandwas laterredesigned

for pitchtrackingoptimisation.This versionhassuccessfullypreventeda recurrenceof landing

PIOsinceits introductionmorethantwentyyearsago.

Criteria evolution

The concept of the synchronous pure gain pilot model became a powerful tool in the discovery of

solutions to high order PIO and design criteria to prevent it. Though the pilot actions were later

found to vary from the pure attitude-related gain model, often with highly non-linear behaviour,

458
the fundamentalpilot actionsare always tightly synchronisedto componentsof the attitude

response.The policy of dealingwith safety-relatedPIO as a specificallylocalisedproblem of

attitudedynamicscompletein itself, separatelyfrom considerationsof generalhandlingqualities,

hasprovedto becorrectandhasledto theauthor'ssuccessfuldesigncriteria.

Theavailabilityafter1978of theLAHOS data,Reference5,enabledthedevelopmentof the

preliminarydesigncriteriondiscussedin Reference4. Thiswasbasedonthenominalstick

pumpingamplitudeandtheattenuationof theattituderesponsebetweenthefrequenciesat 120

degrees(theauthor'sownearly versionof bandwidth)and180degreesphaselag.Thefirst factor

is directlyrelatedto thePIOfrequencyat 180degreeslag,andfavoursahighfrequencyvalue.

Thesecondfactorwasagainmarginof asort,butdid notexplicitly definetheabsolutePIOgain.

TheJaguarFBW demonstrator,designedto thisandother"Gibsoncriteria",beganflight testsin

1981with ahighdegreeof confidencethatthisPIOproblemwouldnotoccur,justified in the

eventasit neverdid.Thismayhavebeenthefirst aircraftcontrolsystemspecificallydesignedto

preventPIOfrom theoutset.

Continuedanalysisof theLAHOS dataresultedin amorecoherentandreadilyidentifiablesetof

parametersenablingapositiveapproachto eliminationof PIOby design.Figure4 (from a 1986

paperandgivenin Reference3) showstheessentialdifferencesbetween"low orderlike"responseswith nosafety-relatedPIOtendencyand"highorder-like"responseswith severe

PIO tendencies.Note that these terms are not usefully related to the actual order of the flight

control system. The most severe LAHOS PIO examples were generated by the addition of a single

lag pre-filter to conventional dynamics, while it is perfectly possible for a 60th order FCS to show

a low order-like response in the critical PIO region. Design criteria based on these observations

utilised the phase rate (similar to phase delay but localised to the 180 degree lag PIO frequency)

and the PIO frequency as shown in the figure, with a maximum permitted PIO gain of one sixth

of a degree per pound of stick force. These criteria, used in the design of the EAP demonstrator,

gave even greater confidence that the PIO problem was defeated. This was again justified by its

extremely successful 1986 to 1991 flight program in which no PIO occurred.

These criteria were incorporated the formal handling qualities specification for the Eurofighter,

which is showing all the excellent handling qualities of the closely related EAP. The design needs

of the fixed gain control mode that was used for a small number of initial flights made it necessary

to identify handling limits that were acceptable and safe rather than excellent, since naturally this

mode could not be optimised for all speeds, especially at touch down. This resulted in further

analysis by the author in 1993 of the LAHOS data to identify PIO gain limits to better quantify

Level 2 and Level 3 PIO effects, and the phase rate metric was modified to the average phase rate

(exactly the same as phase delay but expressed in different units) as a more accurate measure of

high order lag effects. These are shown in Figure 5. (Despite the limitations of the fixed gain

mode, the approach and landing qualities were still very satisfactory).

Some interpretation is necessary in the meaning of the gain limits, as it can be the case that a

response might be classed as Level 2 by its phase rate and frequency, but as Level 1 or Level 3 by

the gain criterion. The author would interpret the gain as signifying better or worse PIO

characteristics, so that any oscillation would be unlikely to diverge with a Level 1 gain but would

probably be divergent with a Level 3 gain. The response should still be classed as Level 2 in the

first case but must be downgraded to Level 3 in the second case.

The author's adoption of "Level" boundaries in design criteria carries no official status, but

reflects only his own analysis of the experimental data based on pilot comments and ratings

according to the "Level" concept.

459
Applicability of Figure 5

The criteria boundaries represent an analysis of a range of response dynamics that is relatively

small compared with the numbers of PIO events that have actually occurred. Many of the

configurations were flown only once by only one pilot, and the opinion rating attached to it might

not be repeated exactly by other pilots. Other configurations might have led eventually to a PIO

given enough exposure to more pilots and more difficult flight conditions. There is a considerable

"grey area" in deciding whether an oscillation should be called a PIO or pilot over-control

resulting from unfamiliarity or insufficient adaptation. It is unlikely that exact boundaries of Level

1, Level 2 and Level 3 PIO qualities could ever be precisely delineated for all examples of high

order PIO.

With three different parameters to be assessed, one of them potentially requiring some

interpretation, it cannot be claimed that this criteria set is guaranteed to quantify with absolute

accuracy the pilot rating of the PIO tendencies of past configurations. What is certain is that the

further outside the Level 1 limit boundaries that the response of a new design penetrates, the

worse its PIO tendencies will be. On the other hand, responses just within the Level 1 limits in all

respects are unlikely to experience significant high order PIO, but they still possess undesirable

residual high order characteristics. The classical aircraft of old without power control actuation

would plot far out of sight to the right on the bottom edge of the phase rate figure, with a response

gain equally far out of sight downwards on the gain plot. Between this ideal extreme and the

practical reality lies a range of increasing high order effects that will eventually lead to PIO

tendencies. Except for unavoidable actuation dynamics, these effects are entirely artefacts of, and

therefore under the control of, the control law designer.

It will be recalled that the definition of Level 1 includes the Cooper-Harper 3 pilot rating with

"some mildly unpleasant deficiencies". A good designer should not simply be content to obtain

the minimum standard just within the Level 1 limits. The designer should set handling qualities

aims equivalent to CHR 2, or better still, CHR 1 which is "excellent, highly desirable". The

concept of an optimum design aim for handling qualities designated Level 1" (Level 1 star) was

used in the EAP control law design guidelines. By illustrating factors that have been associated

with PIO ranging from severe to mild or none at all, the Figure 5 criteria point to the response

dynamics to be avoided by the maximum possible margin to ensure the absence of PIO.

The following Level 1" limits were recommended for linear response design:

• Maximum average phase rate of 50 deg/Hz, equal to a phase delay of 0.07 seconds.

• Minimum attitude PIO frequency of 1.0 Hz.

• Maximum attitude to stick force gain of -20 dB or 0-1 deg/lb at the PIO frequency.

• Maximum attitude acceleration lag of 0.18 seconds in the time response.

(These numbers apply for typical combat aircraft and control inceptors. For other types such as

transport aircraft, similar principles but different numbers may be expected.)

Figure 6 revisits the Tornado configurations, which were rectified without benefit of any proven

criteria, to compare them with the final version in Figure 5. It supports the author's inference that

the first and second pre-filter configurations were not sufficiently different dynamically. The

reliance placed at the time on improving the PIO gain value as a major factor in the solution is

confirmed by the gain criterion which correctly indicates their relative handling. Although the

production version did resolve the PIO problem, it would not pass the later design processes

which led to Level 1" anti-PIO qualities in the EAP for example.

Figure 7 compares the stick pumping at touchdown of the Tornado second pre-filter version in the

incipient PIO incident and the EAP on an early flight touchdown. The sloppy, low frequency and

460
largeamplitudepumpingof theTornadowith about-4--10lbsof stick forceand-4-1_ inches of stick

input compares dramatically with the classically rapid, small amplitude pumping of the EAP with

about 2 lbs of stick force and __ inch stick input, both cases close to the expected frequencies and

producing slightly more than the Bihrle value of pitch acceleration. The high degree of control that

can be exercised by designers over this crucial area of pilot activity is thus clearly demonstrated.

Accounting for actuator saturation

Although the Tornado landing PIO diverged into the non-linear regime of actuator rate limiting, it

was resolved by linear control law modifications. During later development of the "bolt on"

incidence limiting system, actuator non-linearity became a major issue. Linear analysis in the

design stage showed some acceptable reduction in phase margins from the healthy 55 degrees of

the CSAS, and simulation, non-linear modelling and rig tests cleared the system for flight. After

some 40 flights, a very large amplitude self-sustaining oscillation occurred at about 300 knots.

A quasi-linear actuator response model was derived from matching rig tests. Figure 8 shows the

very rapid loss of phase once full rate saturation commenced, typical af acceleration limiting

(Reference 6). This was used to calculate the aircraft attitude dynamics shown in Figure 8. The

dominant feature is the "explosive" growth in the PIO gain as the control inputs become larger. As

the actuator demand doubles from __.7.5 degrees of tailplane to _+15 degrees, the amplitude ratio

quadruples giving eight times the response for twice the stick input. A new non-linear model of

the actuator was also developed with an excellent match to the rig results for all demand

amplitudes. With this model the event could be replicated exactly by analysis. This enabled the

correct design modifications to be developed which effectively linearised the large amplitude

response dynamics, not merely by reducing the phase lag due to rate saturation but by virtually

preventing the occurrence of the saturation altogether.

The most significant factor was found to be the actuator acceleration limiting. The oscillation event

could not be replicated analytically using only the actuator rate limit. This is not usually discussed

in the literature, but it is obvious that the pure saw-tooth waveform often presented as actuator rate

limiting cannot occur in practice. The finite time it takes for the main control valve to be moved

from one end to the other of its stroke represents the acceleration limit. The Tornado tail actuator

control valves were driven by an integrated quadruplex actuator, and though fast it adversely

affected the saturated large amplitude response dynamics. While most fly by wire actuators have

servo drives with much higher bandwidth and rate, the effect of the acceleration limit is always

present and must be included in the actuator modelling for any serious design analysis of large

amplitude PIO resistance.

However, the best means of preventing problems is to provide sufficiently high rates and to

ensure that the forward path command gain at higher frequencies is not unnecessarily large. If the

linear design is also sufficiently low order-like, then the dynamics at the PIO frequency may

change gradually as the input amplitude increases but will not show any sudden and large changes

to trigger a PIO.

Ideally, the rates would be chosen to ensure that the actuation remains unsaturated at frequencies

up to the PIO value using the maximum possible pilot inceptor amplitude. The use of design

inputs smaller than this ignores PIO history. Unfortunately the rates will probably need to be

chosen before the control law design is sufficiently developed to ensure this at critical flight

conditions. A rate sufficient to reach full deflection from neutral in 0-2 seconds permits a full

cycle of maximum amplitude oscillatory control travel while fully rate saturated in 0"8 seconds

(i.e. 1.25 Hz) if there is no serious acceleration limiting. It is hard to imagine that this would not

be sufficient when coupled with proper demand attenuation at PIO frequencies. For lower rates

this attenuation can be adjusted to suit.

461
Thechoiceof desirablemaximumratescanbeconfusedby misunderstandingtheimplicationof

theunitsofrate.High numberstendto alarmmanagement.The importantparameteris how long

it takes for a control to be applied. If a minimum time of 0.2 secondsis desired, the

correspondingrateforroll controlby a differentialtailplanesystemof _+5degreesauthorityis 25

deg/sec(althoughthis would beinadequatefor thetailplane'ssymmetricalpitchcontrolfunction

with perhapsatotal travelof +15 degrees). For a spoiler system with 50 degrees deflection, the

equivalent rate is 250 deg/sec. Allowing for the differing control surface sizes and hinge

moments, the hydraulic power requirements would be roughly similar despite the 10 to 1 range of

angular rates. It is important to get over the message that high rate capability does not mean that

pilots will sit there thrashing the controls at maximum rate for long periods, therefore requiring

large hydraulic power and flow capability. It is only necessary to provide sufficient accumulator

capacity to allow one or two large transient inputs followed by a short dwell in which time the

accumulator can be recharged. It is lack of transient rate capability that can lead a pilot into a

saturated PIO.

Such a provision has been made on the Jaguar FBW, EAP and Eurofighter with actuator rates of

up to 100 degrees per second. Because of their high instability levels, these aircraft could not

tolerate significant rate saturation in the pitch controls. The rudder control rate was also critical,

since its heavy usage to minimise sideslip in providing "feet off" co-ordinated rolling can require

high rates to prevent loss of control in carefree gross combat manoeuvres involving full pitch and

roll inputs in any combination including simultaneously. A second line of defence is to place

software rate limits of a lesser value on the actuator inputs, e.g. 80 degrees per second, so that the

actuators never reach a hard limit. A third defence is to place software rate limits on the inceptor

output signals so that the actuator input rate limits are not invoked or at least are invoked only very

briefly. Inceptor signal rate limiting, being series or open loop, has been found to be tolerated

more readily than closed loop saturation at the actuators. None of these aircraft has shown the

slightest tendency to Type 2 or Type 3 saturation effects in flight.

Designing and testing for good handling

While the thrust of this paper has been the prevention of safety-related PIO, it goes without saying

that the provision of good handling qualities is a necessary precursor. This includes the

prevention of pitch oversensitivity and non-safety-related "low order" PIO such as pitch bobble or

the "PIO syndrome" effect due to excessive attitude dropback or an excessive Bode plot shelf

width. These can easily be dealt with by use of the methodologies described in Reference 3, for

example. Again the designer should aim for "Level 1"" qualities, so that inevitable shortfalls in

some areas will still provide Level 1 handling. Generally this aim can be achieved by a K/s-like

behaviour below the bandwidth frequency, but this must be applied to the appropriate response.

Although control of an aircraft invokes both attitude and flight path, excellent results have been

obtained by optimising the attitude and accepting the fall-out flight path response. This can be

taken only so far, however. The latter may well acquire non-classical features such as "g creep"

and this must always be assessed for acceptability. Flight path control must take precedence in the

landing task, for example, where path control PIO is always a possibility even with classical

response dynamics. Here it is also possible to apply the desired K/s-like dynamics to the HUD in

the form of a quickened climb-dive or velocity vector symbol, giving very precise flight path

predictability and touch down control.

Generally, the faster and higher an aircraft flies, the more dominant the control of flight path

becomes. More strictly, it is control of angle of attack rather than pitch rate that becomes more

important. This is because the steady pitch rate in manoeuvres becomes small relative to the angle

of attack required, which takes too long to acquire initially at anything like the steady pitch rate

462
value.Substantialpitchrateovershootand attitudedropback ratios then become necessary. An

extreme example, discussed (with very approximate data) in Reference 3, is the YF-12 in cruise at

Mach 3 or about one kilometre per second, and hence with extremely low pitch rates per g. Figure

9 shows a time response sketch indicating a good K/s-like path response but an attitude dropback

ratio of 5 and pitch rate overshoot ratio of 6, which are very large by normal standards.

Although such attitude parameters would be highly unsatisfactory in the majority of normal flight

conditions, here their effects are rather insignificant. The normal acceleration increment of about

0.1 lg used to acquire an attitude change of 0"3 degrees for a 1000 foot per minute climb in a

height change manoeuvre required a steady pitch rate of only about 0.07 degrees per second.

Hence the physical dropback and peak pitch rate were about 0.35 degrees and 0.4 degrees per

second. A K/s-like attitude response could be enforced, say by a lag-lead command prefilter, but

the result would be an impossibly long hang-off or g creep as shown in the second sketch.

Despite excellent attitude control, the flight path angle response is made so sluggish that a slow

overdriving PIO would be the most likely outcome of any attempt to acquire a constant altitude or

climb angle. Whether this is truly safety-related is not clear, but it would certainly give a

supersonic airliner captain a hard time with hand flying.

By the start of pre-flight clearance testing, all traces of serious PIO should have been removed by

rigorous design and analysis employing up to maximum amplitude inputs as noted earlier. Even

though this may not represent normal realistic control usage (though it is normal for truly carefree

handling aircraft, where anything goes), a control system unable to withstand this has not been

properly designed. A piloted simulation search for PIO triggers may well be carried out, but

failure to find a trigger task may only mean that the right one has not been thought of. A PIO will

always occur, eventually, if the response dynamics permit it. PIO cannot occur if it has been

designed out of the system, a possibility that has been demonstrated now on several fly by wire

aircraft. A fixed base simulation is certainly capable of showing that Type 2 or Type 3 PIO

characteristics are not present, provided that the control system dynamics are very accurately

modelled from theoretical analysis and rig tests.

After the Tornado, flight testing for PIO at Warton has been confined to a few high pilot gain

precision tasks. One was synthetic HUD target tracking, which showed up a small lateral tracking

oscillation on the EAP caused by a feature introduced to optimise rapid turn entry co-ordination.

On the Jaguar FBW, flight refuelling trials were done at the end of its programme in its most

unstable configuration, without specific pre-task tests but with knowledge of excellent formation

qualities and absolute confidence by then in its freedom from PIO. Eight dry contacts were made

showing very easy control. On Eurofighter, tests of very close formation flying were made

behind a Tornado prior to actual contacts with a Victor tanker. The refuelling task was found to be

an order of magnitude easier than with previous conventional aircraft, and in fact Cooper/Harper

ratings of 1 and 2 were given. Very aggressive pitch tracking has shown an extremely stable

tracking platform. Flight testing for safety-related landing PIO has not been seen as either practical

or necessary given the intense scrutiny applied to the design and pre-flight testing.

Final comments

To design a control system and only then to test it for PIO is a very high risk strategy. To ensure

freedom from PIO, it is essential to plan its absence from the very beginning, starting with a

properly constructed and thought out control law layout, maintaining a highly visible block

diagram on which all paths can be followed and their effects understood, and considering the

impact on possible PIO of the system hardware and of every change to the control laws.

Reference 7, an excellent review of the past PIO problem initiated after the YF-22 PIO in 1992,

recommends a change in paradigm from "Proceed unless a PIO problem is proven to exist" to

463
"Proceedonly when resistanceto PIO is proven". It will be obvious that this author wholeheartedlyconcurs.

Theessenceof safety-relatedPIO preventionby designis simply stated:thePIO frequencycannot

betoo high, thePIO gaincannotbetoo low, thephasedelaycannotbetoo small, andthelarge

amplituderesponsecannotbelinearisedtoomuch.

References

1) Anon, 1991, "AGARD WG.17 - Handling Qualities of Unstable Highly Augmented

Aircraft", AGARD AR-279

2) Gibson, J. C., 1978, "Flying Qualities and the Fly-by-Wire Aeroplane", in AGARD-CP-260,

Stability and Control

3) Gibson, J. C., 1999, "Development of a Methodology for Excellence in Handling Qualities

Design for Fly by Wire Aircraft", Delft University Press, ISBN 90-407-1842-3

4) Gibson, J. C., 1982, "Piloted Handling Qualities Design Criteria for High Order Flight

Control Systems", in AGARD-CP-333, Criteria for Handling Qualities of Military Aircraft

5) Smith, R. E., 1978, "Effects of Control System Dynamics on Fighter Approach and Landing

Longitudinal Flying Qualities", AFFDL-TR-78-122

6) Fielding, C., 1984, "Nonlinear Analysis and Modelling during the Development of an

Aircraft Incidence Limiting System", Chapter 9 in Nonlinear System Design, ed. Billings,

Gray and Owens, IEE, ISBN 0-86341-019-7

7) Flynn, W.A. and Lee, R.E., 1995, "An Investigation of Pilot-Induced Oscillation Phenomena

in Digital Flight Control systems", in AGARD-CP-560, Active Control Technology: Applications

and Lessons Learned.

464
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465
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/ J_.o ( boundaries (Ref. 4)

h. #-"

Safety-related "high order = type PIO region

Figure 2 Frequency response qualities illustrated by non-parametric shape

466
ow_)I 8,

[degAb]

4

..4

/

/ -12

}0

_1/ •" • • • • • • •"" • I Stversion pre-filter

_/" - - - - - - 2nd version pre-filter

P' 3rd version pre-filter

I

m m -.- Unaugmented

I

Version

Time to transient

acceleration peak

PIO frequency

Bandwidth

I

_o(-lzo) Hz

I st 0-245

2nd 0-245

3rd 0-26

Unaug. 0-145

ro(-]so)Hz t,}sec

0-43 0-35

0"455 0"3

0"535 0"22

0-44 0"12

Pre-filter characteristics

(1 + 0-1 57s)/(1 + 0-47s)

(1 + 0"133S)/(1 + 0"454s)(1 + 0"0625s)

(1-0)

(1-0)

Figu re 3 Tornado pitch attitude responses at landing: solution to PIO

by development of the command pre-fllter.

The unaugmented and third version pre-filtered dynamics are PIO-free.

467
[d_Abl /

-220 -lt_n -140 -100 9_'[deg]-60

l -%1

"1

Pilot gain for PIO r'_ .-,"""_ "_'_

" ---_.'_ 5-4 0"5dlr ---o--''" High order response type:

o.5 Jlr-_ _PIO likely because of

,, IF I_1 t_ i - Low PIO frequency

1L i _ 1 _" - High aircraft gain

, ' _ 1 _ - Large phase rate i F_;7

Pilot gain for PIO -24.L Jp1 _ y Low order response type:

is 25 Ib/deg I _ PIO unlikely because of

_- High PIO frequency

t_ - Low aircraft gain

1"st_ ' - Low phase rate

hase rate: the local slope of phase lag

vs. frequency at the -180 ° phase angle point

Phase

rate

[deg/Hz]

t

Severe:,,_O t. .... I_"_e 1"3

2°°t V ....

/ _ M°clerate P'O "_=-

0 0-4 0"8 1 "2 1'6 2"0

£0(-180) [Hz]

Figure 4 PIO tendency indicators and design guidelines

derived from LAHOS etc.

468
le (0)

-4 _,_[deg] \

/

Exampleof_timum respon_j

to resist PIO _-",,,._

,f

_ Amplitude boundaries:

--y,V_'-/" Upper limits of attitude

-20_"" frequency response in

"L! _ the PIO region

_f_Level 1° upper limit

• for optimum design

(a) PIO gain limit criterion

I Phase[deg]

phase angJ

I= _-oJ [Hz]

X_ i °J(-ls°)

2 _(_ I_)

Average phase rate

= {- A phase angle + ¢u_-lso_} 0eg/Hz

Phase delay z'p

-_ I Average phase rate _- 720 }

(b) Definition of average phase rate

25C

20C

Average

phase

rate 15(]

[deg/Hz]

10C

50

f

f

//'-

0

0 0-2 0-4 0-6 0-8

L3

L2

L1

deg/Hz (rp)

195 (0.27)

145 (0.20)

85 (o 2)

Opt!mum L_vel 1"'--''_

des, gn region

1-0 1-2 1-4

w(- _so) tHz]

(c) Phase rate and frequency criterion

Fkjure5 Final development of PIO criteria (1993)

1. Level 1, 2 and 3 boundaries represent historical data.

2. Undesirable residual high order characteristics exist within the

Level 1 region near the low frequency boundary limit.

3. Best design practice for freedom from linear high order PIO requires

the more stringent Level 1° gain, phase rate and frequency limits.

469
1 1,8I

Ne_b] 4 t _...._/

I

_°.,_..,..J,""-,,t %0/ ,l-// -,:6o ..,"-,-_'..,_'-,,_oo .-/,..,.. C'_,,.,.._,,.-?-,; ¢[._

_,,,-_r_,/l. ......-- - - - - 2nd version pre-filter

L1/_ 3rd version pre-filter

4'

--- --- --- Unaugmented I

25t:

20C

Average

Phase

Rate 15£

[deg_Hz)

10(;

5O

1st pre-f_te_/,s, _ L3

/_ 1_2nd pre4ilter L2

0

0 0-2

I

deg/Hz

195

145

_-3rdwe-filter L1

, I' ' I I

I _ Unaugmented ii kl*

I I I

0"4 0-6 0-8 1-0 1-2 1-4

_-]_O_[Hz]

Figure 6 Tornado viewed in retrospect against author's later criteria

Note: although the 3rd pre-filterjust satisfies the criterion and

has prevented PIO for 20 years, it would not have been

accepted as a new design by subsequent criteria.

470
Pitch attitude [_g] _8_ +0-48°

..........

_ ?i_ X, ..iat.i .... ld__.F: .......... - N..... __..... "L"":I:>'____' ': ..........

....... _.......... II:-::I':' ' , .: .I :.::" .. .................. "42e,_ " _

.-_51-+,:o ' I .......... .,11.... \ ......

i ........:j.,, stickforc_IN],_--,, ..... _' ......................... '...........

- -_--.i I s_ck pumpincj . _nc_entP,O Tornado

- "- Pgotattempts to - (2nd control law)

freeze the stick

_i 'o !

.............b, _,...................................... :.i ......................."

PITCH

S r RATE

i' : ..1:7

-s _ ..:::,,_:.:i::e....._...o:._.r._..-..::..::.::::,::I..._:::..:.:l:;..__

" -' ; o ,= "t.- ; '

ITICK

PUI'IP | NI I- --t

o-,_r i....... i g'-J'-'--

• I .......... ._=K ........... -_- .......... r,,._.....................

o '_,--_--_. _.:-_.._a.__ 'g_ , ---

_ ./ }.,,,v,., "x_." _.

"" ,I,IN

EAP

Figure 7 Effect of design process on stick pumping and

associated PiO resistance

471
oI ....._-_.__... Linear

_.,_d_I :'_'__ deg __ +_3deg

_IL_-_';"_i-_i i_.

I

, _.,.:.--...-..:_--_ _'.._\_iii; re_ de,i_ed,,om

90_ rigtests

0-1 1.0 5"0

Ir--_lOB /

Id_Ab] 4

°,,.o,..._, :;;--..: :.p>- -"_'_I_Iial -ratelillitonset

Unear response

Rate saturated

response

Figure 8 Significant non-linear actuation effects on PIO characteristics

472
Pitch

attitude

Flight

path

angle.

Attitude dropback ratio = 5"5 sac. 1=,_¢_6_/

Pitch rate overshoot ratio = 6 _ .,-,

Path ttme delay = 0-5 sec./- ,,,-'_,_e

0 _ 1_) 1_5 secs. 2_)

Nominal YF-12 time response

at Mach 3 cruise

I

-5

Flight [/ / ._,_@_

angle i_.:_c,_]_.,__ v,i_

6 lb secs

/

Short period roughly approximated by:

¢Osp= 0-5 rad/sec

_ = 1"3

Path delay = 5-2 sec.

YF-12 with pre-filter =

1 +0-8s

1 +5-5s

Figure 9 Sluggish PIO-prone flight path response caused by

inappropriate pitch attitude optimisation

473
q
98-4333

PILOT-INDUCED OSCILLATION PREDICTION

WITH THREE LEVELS OF

SIMULATION MOTION DISPLACEMENT

Jeffery A. Schroedcr', William W.Y. Chung*, Duc T. Tran"

NASA Ames Research Center

Moffett Field, California

Soren Laforce, Norman J. Bengford

SYRE Logicon

Moffett Field, California

Abstract Fto,,F_°,, Ftee long., lateral stick and pedal force, lbs

litd

Simulator motion platform characteristics were K

examined to determine if the amount of motion affects KH

pilot-induced oscillation (Pit) prediction. Five test Ko

pilots evaluated how susceptible 18 different sets of Ls,t

pitch dynamics were to Pits with three different levels M_

of simulation motion platform displacement: large, Ns,_

small, and none. The pitch dynamics were those of a n

previous in-flight experiment, some of which elicited p

Pits. These in-flight results served as truth data for the s

simulation. As such, the in-flight experiment was To_,Te2

replicated as much as possible. Objective and [3

subjective data were collected and analy7ed. With large fie

motion, PIt and handling qualities ratings matched the 5c_

flight data more closely than did small motion or no 5_t

motion. Also, regardless of the aircraft dynamics, large 5eaick

motion increased pilot confidence in assigning handling 51o_,_,.,8 _

qualities ratings, reduced safety pilot trips, and lowered

touchdown velocities. While both large and small

motion Ixovided a pitch rate cue of high fidelity, only _o,

large motion presented the pilot with a high fidelity _p, _p

verticalaccelerationcue. _l, _2

a,b,c

amo_et

amotiou

F(x,y)

o,_

ptefilter zeros and poles, tad/see ¢,, ¢,

model acceleration, ft/sec 2,tad/see 2 coa,

motion system commanded acceleration, corot*

ft/sec 2,rad/sec 2

variance ratio with x and y degrees of ¢e_, cop

facedom

touchdown vertical velocity, ft/scc

control system prefilter gain

motion system filter high-freq gain

control system gearing, deg/in

lateral control sensitivity, 1/sec2/in

elevator control sensitivity, |/See 2

directional control sensitivity, 1/seca/in

number of points in each mean

probability that effects are random

Laplace transform variable, rad/sec

pitch-to-elevator zero time constants, sec

sideslip angle, deg

elevator deflection, deg

commanded elevator, deg

filtered commanded elevator, deg

commanded elevator from stick, deg

longitudinal, lateral stick and pedal

deflection, in

Dutch roll damping ratio

motion filter damping ratio

phugoid and short period damping ratio

control system prefilter damping ratios

complex zero damping ratio in bank-toaileron transfer function

pitch and roll angles, deg

roll and spiral mode time constants, sec

Dutch roll natural frequency, rad/sec

motion system filter natural frequency

rad/sec

phugoid and short period natural

tad/see

*AerospaceEngineer.SeniorMemberAIAA.

t AerospaceEngineer. MemberAIAA.

Copyright© 1998 by the AmericanInstituteof Aeronauticsand

Astronautics,Inc. No copyrightis asserted in the United States

underTide17.U.S. Code. TheU.S. Govemn_nthas a royalty-free

license toexercise allrightsunderthe copyrightclaimed hereinfor

GovernmentalProposes. All other rights are reserved by the

copyrightowner.

1

American Institute of Aeronautics and Astronautics

475
control system prefiiter natural

frequencies, rad/sec

complex zero natural freq. in bank-toaileron transfer function, rad/sec

Introduction

from the visual field-of-view, which corresponded to an

altitude of 100 ft.

Figure 1 shows the desired touchdown point, which

was the near-left comer of the 1000-ft fixed distance

marker located to the right of centerline. This desinxl

touchdown point matched the flight-test study. Table 1

gives the performance standards for the task.

Ground simulation has not been very successful at

predicting subsequent in-flight pilot-induced oscillations

(PIOs). A recent study recommended that "validating

simulation details, protocols, and tasks and collecting

and correlating them with flight test results should be

given high priority" to improve this simulation

we_rdless.

With two fixed-base simulators of different

capabilities, Ref. 2 evaluated the longitudinal PIO

tendencies of configurations tested in a PIO flight test

study) The simulation results followed the general

t_end of the in-flight data; however, the worst in-flight

configurations were not as severe on either fixed-base

simulator.

The purpose of this study was to determine what

effect simulator platform motion has on predicting

PIOs. Here. three simulator platform motion

characteristics were examined: large, small, and no

motion. Five pilots flew a landing task with 18

different sets of longitudinal dynamics with each motion

configuration. Both pilot-vehicle performance end

subjective data were taken and compared with the

previous in-flight study)

Apparatus and Tests

Task

The in-flight task was replicated as much as

possible: Pilots started at 135 knots and 1.5 nmi from

the runway and flew three visual approaches to full

touchdown with each configuration. One approach was

straight-in, and one each started with a 150-ft left or

right lateral offset from the touchdown point. During

the approach, pilots were instructed to maintain

constant speed and remain on the glidepath (-2.5 degs)

and localizer. Deviations were indicated on heM-down

instruments. At the start of the run, the aircraft was

placed 1/2 dot off the desired localizer and glideslope.

For the left and right offsets, pilots held that offset

until an automated voice instructed the pilot to

"correct." The pilot then maneuvered the aircraft to land

on the desired touchdown point. The "correct" command occtmed when the runway overrun disappeared

2_

mooft 2,2 Desired

touchdown pt

Figure 1 - Landing task

Table 1- Task performance standards

PIOs None None

Longitudinal

touchdown +/- 250 ft +/- 500 ft

¢lTOr

Lateral

touchdown +/- 5 ft +/- 25 ft

error

Approach +/- 5 kts -5/+10 kts

Math model

Longitudinal configurations. A linear stability

derivative model 4 generated the aerodynamic forces and

moments on the aircraft. Bare airframe derivatives wee

combined from several sources. TM Response feedbacks

of angle-of-attack and pitch rate to the elevator were

used to simulate the different pitch configurations,

given below, which mimics the NT-33 variable

stability aircraft,s Figure 2 shows the dynamic blocks

of the pitch axis dynamics.

The simulation centerstick dynamics were measured

as:

0.125(222 )

_n (s) s2 + 2(0.7)(22)s + 222

2

AmericanInstitute of Aeronautics andAstronautics

476
Thesedynamicsareslowerthanthe25tad/seestick

longitudinal natural frequency stated in Refs. 3 and 7

due to force-feel system limitations of this simulator

cockpit. The ergonomics of the stick matched Ref. 7.

I 01'°rl

I

Figure 2 - Longitudinal block diagram

Fourteen prefilters were simulated as in the inflight experiment. These prefiiters consisted of first,

second, and fourth-order linear filters. These fliers are

of the form below, and Table 2 gives their values:

__Set-tO(s) = K(s+ a)

6ea_, s+b

Sco.,, (s) = K

Be.,,_ s + c

5ec K

-_-=2._- (s)ffis2+ 2_lcols + col2

0emr.t

_ec_. (s) = K

_es_.t (52 + 2_Icol s+O}2)(s2 +2_2_2 s+(l}2)

Filer

Table 2 - Control system prefllters

K a b c _ o_ G o,

B 3.0 3.3 10 .....

D 0.5 20 10 .....

1 1.0 .......

2 10 - - 10 ....

3 4.0 - - 4 ....

5 1.0 - - 1 ....

6 162 - - - 0.7 16 - -

7 122 - - - 0.7 12 - -

8 92 - - - 0.7 9 - -

9 62 - - - 0.7 6 - -

10 42 - - - 0.7 4 - -

11 16' - - - 0.93 16 0.38 16

12 22 - - - 0.7 2 - -

13 32 - - - 0.7 3 - -

Commanded elevator deflection was the sum of the

prefiiter output and the feedbacks of angle-of-attack and

pitch rate. The elevator actuator dynamics were modeled

as a second-order filter with the NT-33 rate and position

limits. 7 In the linear range, the actuator dynamics are:

_C(s) = 752

5ec s2 + 2(0.7)(75)s + 752

Four sets of aircraft dynamics were evaluated. The

differences among the dynamics were effectively in the

short-period mode. The pitch-to-elevator transfer

function had the following form:

0 MS, (s + 1/T0t )(s + 1/To2 )

--(S)_" 2 2 2 2

Se (S +2_pcopS+COp)(S +2_spcospS+cosp)

Table 3 gives the parameters for the above transfer

function. For all configurations, M_,=-3.3 l/sec 2.

Table 3 - Aircraft dynamics

AJC Tm T_, _ COp _p COn

2 12 1.4 0.15 0.17 0.64 2.4

3 12 1.4 0.17 0.16 1.0 4.1

4 12 1.4 0.16 0.16 0.74 3.0

5 12 1.4 0.16 0.15 0.68 1.7

The remaining parameter to be specified is the gearing between the elevator command from the stick and

the longitudinal stick position. For the 18 tested configurations, which represent combinations of the aircraft

dynamics and prefilters, the gearings are listed in Table

4. As an example, for configuration 2-B, the "2" corresponds to the values in Table 3 and the "[3"

corresponds to the values in Table 2.

Subsequent to the experiment's start, information

from the Ref. 2 authors indicated that the Table 4

gearings may have been 70% higher than in the flight

test. To evaluate the effect of different gearings on the

results, a mini-experiment was run using the Ref. 2

gearings with configurations 3-1, 3-D, and 3-12.

Differences between gearings were less than or equal to

one handling qualities and pilot-induced oscillation

point.

Each of the 18 configurations was verified by

performing frequency sweeps on each and overplotting

the result against the analytical pitch-rate-to-stickdeflection transfer functions.

3

American Institute of Aeronautics and Astronautics

477
Table 4 - Gearings 

Config & Config & 

2-B -2.94 3-8 -7.29 

2-1 -2.94 3-12 -7.29 

2-5 -4.33 3-13 -7.29 

2-7 -2.94 4-1 -3.46 

2-8 -2.94 4-2 -3.46 

3-D -8.65 5-1 -1.73 

3-1 -7.29 5-9 -1.73 

3-3 -7.29 5-10 -1.73 

3-6 -7.29 5-11 -1.73 

The engine model consisted of a first-order transfer 

function from throttle input to thrust output. The time 

constant was nonlinear and depended on RPM.’ 

Using a lateral-directional stability 

derivative model, coefficients were adjusted to achieve 

the following modal and sensitivity characteristics: 

5, =0.3 sec 

5, = 75 sec 

a& =a* =1.3 rad/soc 

L*Iu -0.7 rad/sec2/in 

Ngd =-0.2radIsec2/in 

These characteristics were also verified with figquency 

sweeps. 

Amrosbhere. Dryden turbulence with rms 

magnitudes of 3 ftlsec was used. A vertical lcosine 

gust occurred whcn the aimaft reached an altitude of 100 

fi. The gust had a peak of 12 Wsez and was time scaled 

based on the 6.7 ft chord of the NT-33. 

Evaluation pilots in the NT-33 flight 

study were accompanied by a safety pilot, who ended the 

evaluation and assumed control of the aircraft if a 

potentially hazardous situation occd. If a safety 

pilot assumes control, then questions arise immediately 

on that configuration’s “controllability” from the 

handling qualities point of view. The presence of a 

safety pilot can also add a factor of stress, since another 

set of eyes is watching thc evaluation pilot. 

In this simulation, an automatic safety pilot was 

implemented that assumed control of the simulated 

model when the nosewheel’s vertical speed exmded -8 

Wsec below a center-of-mass height of 12 feet This 

criterion was developed empirically and was well 

received by the pilots. Upon activation, the pilot’s 

controls went dead, a voice said “my airplane.” and the 

math model initiated a go-around. 

WatoE 

on svstem. The NASA Ames Vertical 

Motion Simulator (VMS) was used.’ It is the world’s 

largestdisplacement flight simulator, with capabilities 

shown in Figure 3. The cockpit was oriented for large 

longitudinal travel. The dynamics of the motion 

system were measured during the experiment using 

frequency response testing techniques? These dynamics 

were fit with an equivalent time delay in each axis. 

Software feedforward filters were used to tune the delays 

to achieve a close match among axes. The equivalent 

time delays for the surge, sway, pitch, roll, and yaw 

axes were all 80 msecs, and the heave axis had 110 

mscc of delay. By comparison, delays in the NT-33 

model following control system have been suggested as 

being in the 45-60 msec range. 

Figure 3 - NASA Ames Vertical Motion Simulator 

Visual svstem. The visual scene was rendered with 

an Evans & Sutherland BIG-3000 image generator. 

Three monitors comprised the field of view, as shown 

4 

American Institute of Aeronautics and Astronautics 

418 
in Fig. 4. The visual system had a measured time dehy 

of 80 msec from the pilot's stick position to the visual 

scene. Figure. 5 shows the visual scene with the ahaft near the runway. The nose of the simulated aircrali is at 

the bottom of the field-of-view. Window mullions were 

added (ovd in Figure 5) to repticate the cockpit.' 

Figure 4 - Cockpit field-of-view 

Figure 5 - Simulator cockpit photo 

CQ&L The lateral stick and pedal dynamics wae 

mcasurcd as: 

0.25(16*) 

s2 + 2(0.7)(16)~ + 16* 

aped 0.0 1 67(ZS2) -(s) = 

Fped s2 + 2(0.7)(25)~ + 252 

A bead-up display was video mixed with the visual 

scene. The display includcd a pitch ladda, altitude 

above sea level, airspeaa rate-of-climb, heading, range, 

and a flightpath marker. The flightpath marker 

reprcsmtad centa-of-mass flightpath and used raw data 

only. 

Motion co -ratioThree motion configurations were examined: large. 

small, and no motion. The VMS motion platform 

soffware was modified to implement each 

Thc classical washout motion 

control laws of the VMS wae used for this 

configuration. second-ordn high-pass (washout) filters 

exist between the math model accelerations and the 

commanded motion system accelerations. Thest tiltas 

have the fom: 

In each of the six motion degrees-of-freedom, both Go, 

and a,,, wae adjusted to keep the motion system 

within its displacement limits using motion system 

fidelity criteria suggested initially by Sinacori" axi 

rcvistd and validakd subsequently." Table 5 shows the 

values used. The damping ratio, CmM, was 0.7. In 

addition to these cues. rolYsway coordmh . 'OIL aad 

residual tilt crossfeeds wa-c present in the motion 

Iogic.'2 

Table 5 - Large motion system ~anrm etm 

Axis a,,, 

Pitch 1 .oo 0.20 

Roll 0.40 0.50 

YaW 0.65 0.20 

Longitudinal 0.65 0.40 

Lataal 0.50 0.50 

Vertical 0.80 0.30 

A coordinated-adaptive algorithm, 

used on many of today's hexapods. was employed in the 

small motion configuration."J' This algorithm 

assumed a mathematical model of a hexapod platform 

with Win stroke actuators. Thus, the stroke limiting 

that occurs when commanding several axes was present. 

Euler angles and translational positions of the platform 

were back solved on line from the resulting (ad 

potentially limited) actuator position^.'^ The ma 

angles and positions wae then used to drive the VMS 

platform. 

scombnk high-pass filters wclt d in tbc 

translational axes, while the rotational axes used a firstadp high-pass filter (unlike the Large motion 

configuration). The de filters had a damping 

ratio of 0.7, except for the surge axis, which was 0.8. 

For comparison. Table 6 gives the gains and natural 

5 

American Institute of Aeronautics and Astronautics 
frequencies (or pole locations) for the small motion

filters. The gains listed are the maximum values, as the

coerclinmd-sdapdvealgorithinreduces these values

when the actuators near their travel limits. These gains

were adjustedto use as much of the 60-in actuator

strokeas possible.

Table 6 - Small motion system parameters

Axis K_ (o,,m (or

pole)

Pitch 0.50 0.30 (pole)

Roll 0.25 0.81 (pole)

Yaw 0.70 0.30 (pole)

Longitudinal 0.11 0.67

Lateral 0.45 0.90

Vertical 0.13 0.90

The motion system was tw'ned off in

this configuration.

_jxnmrison with fidefi_ criteri& Figure 6 plots

each axis of the large and small motion configurations

against the validated criteria of Ref. 11. These points

are determined by finding the magnitude and phase of

the respective motion filter evaluated at I rad/sec.

: - large small motion [

Low Fidelity

o _edium

Roll/ Fi_...-,-_

/ / mHighFidelity

[ [ _'[_h lYamw , Pitcll 0

0.0 0.2 0.4 0.6 0.8 1.0

Rotational gain @ 1 rad/sec

D []

Vert. Lat.

CI Low Fideli_..,......_

Long. o _ Medium

/ / High

, , I I o

0.0 0.2 0.4 0.6 0.8 1.0

Translational gain @ 1 rad/sec

Figure 6- Motion fidelity prediction

30

60 Phase

orrof

(deg)

4O

20

80

60 Phase

PAYor

(deg)

4O

_0

In the rotational axes, high motion fidelity is

lmxiicuxi for both pitch and yaw motion with the large

end small motion configuration. Roll motion is low

fidelity in both motion configurations, since the roll

axis was attenuated to minimize the false lateral specific

force cueing during coordinated rolling maneuvers.

In the n'anslafional axes, all of the small motion

cues are predicted to be low fidelity. For large motion,

the fidelity improves, cspecial]y for the vertical axis,

which provides n key cue for this task. This figure

shows the benefit of large motion in fidelity terms.

Pilots

Five expericnce test pilots, _ refined to as A-E,

participated. Pilot A was an FAA test pilot, pilots B-D

were NASA Ames test pilots, and pilot E was a Boeing

test pilot.

Exoerimental procedure

Summarizing the experimental variables, they were:

1. motion configuration (3),

2. aircraft confignration (18)

Thus, each pilot evaluated 54 configurations. Pilots A,

B, and E evaluated each configuration at least twice.

Pilots C and D evaluated each co_surafion only once.

The pilots each n_l the same experimental

briefing. They had no knowledge of the configurations,

which weze ramicmdzed. After flying the task, the

pilots were told of their perfemmnce. Then, they

assigned a handling qualities rating using the CooperHarper scale, Is a Pilot Confidence Factor? 6 and a Pilot

Induced Oscillation Rating (PIOR). 6

Results and Discussion

_.tllxt_tlma

Figure 7 illustrates a classic

divergent PlO that occtmed with Pilot B, configuration

3-12, and large motion. The pilot was nearly on the

longitudinal stick stops. The pilot gave this configuration a Cooper-Harper rating of 8, and a PIO rating of 5.

PlOs of this severity end for this exleackd periecl of

time did not occur for eithor the small or no motion

configurations.

The average frequency of the PIO in Figure 7 is 3.0

ra_sec (the average in-flight PIO fiequency of this

6

American Institute of Aeronautics and Astronautics

480
configuration was 2.2 tad/see). Also shown on the

pitch rate and normal acceleration traces are the motions

that both the large and small motion configurations

would produce for this visual motion.

&0

|o

f

40 5' 10' 'S ' ' ' ' I 2O ea 26 26 4O

-7,S S 1 16 20 26 40

i7 ....

j/22 +!

T_ra {me}

Figure 7 - Example PIO

At the PIO frequency, the large motion

configuration provides 100% of the pitch rote cue, and it

leads the visual scene by only 5 degs of phase angle.

So, the dashed line overlays the sofid line. These

values may be determined by inserting 3 red/see into the

motion system filter discussed earlier with the pitch

axis parameters (Table 5). The small motion configmw

tion, at best, provides 50% of the visual pitch rate and

leads the visual by 6 degs. By motion cueing fidelity

standards, both the large and small motion cues are high

fidelity. I°.n

For the normal acceleration, the large motion

configuration provides 80% of the visual cue and leads

the visual by 3 degs (this value includes the motion

filter sad the additional 30 msec of delay that the

vertical platform lags the visual). But the small motion

configuration provides only 13% of the visual cue end

leads the visual by 20 degs. By motion cueing fidelity

standards, the large motion cue would be high fidelity,

and the small motion cue would be low fidelity. It is

fc_ this important acceleration cue that large motion

provides a simulation benefit, and it is likely the reason

for the superior perfm.anee of the large motion

configuration as discussed later.

T_nrlin_ l_n'fommnce. Lon_tudinal touchdown

position was analyzed using a two-way n:Ixaled

measures analysis of variance (ANOVA). t7 While

statistically significant differences occuned across the

ahcrait configurations (F(17,68)=3.73, p<0.001),

differences among the motion configurations were not

found (p>0.2).

Lateral touchdown position was analyzed, and no

significant diffe_tces were noted among the aircraft

(p>0.4) or motion configurations (p>0.4). Approach

airspeed en_rs wae almost always within the desinxl

pe_ormm:e standa_

During the evaluations, it was noticed that pilots

had difficulty in judging sink rate during the flare-totouchdown as less platform motion was ixesonted.

Indications of this fact were either harder landings or the

safety pilot assuming control for the small and no

motion configurations.

Figure 8 shows the means and standard deviations

of vertical touchdown velocities for each motion

configuration. Each mean is an average of 90 points

(18 configurefions x 5 pilots). The ANOVA on these

data indicated that the motion configuration affected

touchdown velocity indcpeedent of the vehicle

configuration (F(2,8)_36.8, p<0.001)? 7 Aircraft configuration also affectedtouchdown velocity indepeadent

of motion configuration (I:(17,68)=2.93, p<0.001). No

interaction between the motion and vehicle configuratious was present (p>0.3). Thus, touchdown velocity could be modeled as independent functions of the

motion and aircraft configurations:

i_td= f(motion) + g(aircraft)

As more motion was available, pilots were able to

lower the touchdown velocity. A previous limited

experiment with large motion also indicated this effect

when the longitudinal handling qualities were poor, n

however, the results here indicate that large motion

allows lower touchdown velocities regardless of the

configuration.

As Table 1 notes, sink rate at touchdown was not a

performance parameter in this experiment, which was

also the casc in thc Ref. 3 flight experiment. However,

the Ref. 2 simulation expcfimont addai a touchdown

perfonmnco criterion of _ 4 ft/sec for desired paformance and _;8 ft/sec for adequate performance. Had that

been the case here, it is expeet_ that even furtlzr

differences among the motion configurations would

7

American Institute of Aeronautics andAstronautics

481
have occurred. This is because when more platform

motion was added, it compensated for sink gale

perception deficiencies in the visual scene.

-4

-7

n=90

t | t

None Small Large

Motion Configuration

Figure 8 - Touchdown velocities

Figure 9 shows the number of

times the automated safety pilot assumed control versus

the motion configuration. Over 1400 landings wezc

performed, so the safety pilot assumed conUrol in

approximately 10% of the landings. It took control

slightly fewer times with small motion than with no

motion; however, large motion resulted in significantly

fewer safety pilot trips. Many of the safety pilot trips

occurred from the inability to judge sink rate.

Wlfile it was stated earlier that causing the safety

pilot to assume control should raise questions about the

configuration's controllability, this seldom coctamL

Pilots often felt they were still in control. The issue

was thatthe small or no motion configurationsdid not

assist pilots in their estimation of vertical velocity as

did the large motion cues.

Longitudina] stick rms positions

were analyze& Statistical dif[enmces occurred across

aircraft configurations (F(17,68)=7.81, p<0.001), with

configurations 5-10 and 3-12 having the most activity

(0.96 and 0.95 in, respectively). Configurations 2-B

and 3-D had the least activity (0.49 and 0.51 in,

respectively). No significant differences occurred across

the motion configurations (p>0.1).

75

J

5O

"6 25

t |

None Small

Motion Configuration

FigzLm 9 - Safety pilot trips

|

Large

Handlin_ Oualities Ratln_s

Figure 10 is a plot of the in-flight

HQRs 3 versus the simulation HQRs for the large

motion condition. If simulation matched flight, then

all points would lie on the diagonal line. A l-unit

HQR band is plotted about this line, which is often

taken as the range of an acceptable match. Eight of the

18 configurations lie within this 1-unit band. Very

similar t_nds to that of the Ref. 2 fixed-based

simulation _e noted. That is, the best configurations

in flight were slightly worse in simulation, and the

worst configurations in flight were better in simulation.

10 3-13 5-10 ///I

9 2-8)E 2,)1( ,'/,_/.

8 3"1_///, ` ]

_. 7 3.8. .s.9/ " / / / t

5 //

//

4 //

1

I 2 3 4 5 6 7 8 9 10

HQR Large Motion Simulation

Figure I0 - Hight versus large motion HQRs

8

American Institute of Aeronautics and Astronautics

482
Figure 11 shows the in-flight

versus simulation HQRs for small motion. Six of the

18 configurations lie within the 1-unit band, which is a

degradation from the large motion condition. Again,

the same trend on the best and worst configurations

existed as for large motion.

Z"

UQ

M.

a:

O

-i10 • . • X. ._K. . .- #,

3-13 5-10 / J

9 2.8_ _2.5 / /_f ,/,

I/ tt

8 3-12_//// q

7

5"1J)_/ / /

6 / /

. /

4 //_3-3

/ _'2_, _"'_5-1

z o,

2 '_/3-._ _.1

/.," 1

1 2 3 4 5 6 7 8 9 10

HQR Small Motion Slmulatlon

Figure 11 - Right versussmall motion HQRs

Figure 12 shows the in-flight versus

simulation HQRs for no motion. Five of the 18

configurations weze within the l-unit band, which is a

degradation from large motion and small motion.

Again, the same tend on the best and worst

configurations existed as for large andsmall motion.

o)

I/.

CC

O

-r"

10

9

8

7

6

5

4

3

2

1

• . .

2-8)1( 2-6_ / / /

3-1_1(/////

3-8)1(_5-9/////

5-11_////

,f , •

2 3 4 5 6 7 8 9 10

HQR No Motion Simulation

Figu_'c.12 - Flight versus no motion HQRs

Pilot Confidence Factors. Confidence factors of A,

B, and C refer to a pilot's opinion that he can assign a

handling qualities rating with a high, moderate, or

minimum degree of confidence, respectively, s6 Losses

of confidence arise when simulation cues are incomplete

or inadequate. Figure 13 shows that as more motion is

provided, the pilot's confidence in assigning ratings

improves. On average, both the no motion and small

motion configurations caused the pilot to have less than

a moderate degree of confxlence in his rating. With

large motion, that confidence improved to mare than

moderate. This difference was statistically significant

across the motion configurations (F(2,8)=5.82,

p=0.028). Diffesences in this measure wan not

significant across the aircraftconfigurations (p>O.l).

_B

8

_.c

n=90

I I I

None Small Large

Motion Configuration

Figure 13 - Pilot confidence factors

llO..JCatta_

Rgurc 14 compaz_ pilot-induced

oscillation ratings (PIORs) between flight and the large

motion simulation. Sixteen of the 18 configurations

lie inside the +/- I PIOR boundary. Except for four

configurations, the in-flight PIORs were, on average,

higher than the simulation PIORs.

9

American Institute of Aeronautics and Astronautics

483
....

////

5 /

_ ,"

_4 2-8_5-9 _ /

_. 3-8_ /" / j"

rr / $-11_ /

03 /'2-7_y /

i. / / _a./ //

3- -/ . 2-1 /_

1__? ' '

14-1 2 3 4 5

PIOR Large Motion Simulation

Figure 14 - Flight versus large motion PIORs

PIORs for the small motion

configuration are shown in Figure 15. Here, 12

configurations were inside the +/- I PIOR band, which

was the worst performance of the motion configurations. Again, except for four configurations, the inflight PIORs were worse than the simulator PIORs.

_" / 8-e}l( ," / "

,,"

2_ i-B /

13-D 2 3 4 5 6

PIOR Small Motion Simulation

Figure 15- Flight versussmall motion PIORs

The PIORs for no motion are given in

Figure 16. No motion perfcmmedslightly better than

small motion, but worse than large motion. Fourteen

configurations were inside the +/- 1 PIOR band. Still,

except for four configurations, the in-flight PIORs wexe

higher than the no motion PlORs.

6

//

5 /)t_lo

)K3-13 2.5)1(/)1(3-1y / /

_4 2-8)1( /'t/.5-9 / /

E 3-8)1( // / //

E / 5-11 / /

';3,'1/ / //

,X/_"4"2_- _'o r,.1

1 4-1 2 3 4 5

PIOR No Motion Simulation

Figure 16 - Flight versus no motion PIORs

6

A piloted experiment examined the effect of three

levels of platform motion displacement on the ability to

predict pilot-induced oscillations. Objective and subjective measures were examined for large, small, m_dno

platform motion. The small motion condition reproseated the displacement of a conventional hexapod

platform.

Overall, large motion matched fright more closely

than either small or no motion. Specifically,

motion better matched the in-flight pilot-induced

oscillation ratings and the handling qualities ratings

than did small or no motion. In addition, with large

motion, pilots assigned higher confidence factor ratings,

achieved lower touchdown velocities, and caused fewer

safety pilot trips as compeaed to the other motion

configurations. Finally, only with large motion did

markedlydivergent pilot-induced oscillations occ_.

An example illustrated that high fidelity pitch rate

cues were provided by both the large and small motion

configurations. However, only large motion allowed

high fidelity vertical acceleration cues to be presented.

Pilots react strongly to vertical acceleration, and this

fikely contributed to the large motion cenfigurafion

providing the best results.

1. National Research Council Committee on the

Effects of Aircraft-Pilot Coupling on Flight

I0

American Institute of Aeronautics and Astronautics

484
.

.

.

5.

.

7.

.

.

Safety, Aviation Safety and Pilot ControlUnderstanding and Preventing Unfavorable PilotVehicle lnteractiona, National Academy Press,

Washington, D.C., 1997.

Kish, B.A., Leggett, D.B., Nguyen, B.T., Cord,

TJ., Siutz, G.J., "Concepts for Detecting PilotInduced Oscillation Using Manned Simulation,"

AIAA Paper 96-3431, San Diego, CA, 1996.

Bjorkman, E.A., "Flight Test Evaluation of

Techniques to Predict Pilot Induced Oscillations,"

Masters Thesis, AF1T/GAE/AA/86J-1, Air Force

Institute of Technology, Dec., 1986.

McRuer, D., Ashkenas, I., and Graham, D.,

Aircraft Dynamics and Automatic Control,

Princeton University Press, Princeton, NJ, 1973,

pp. 203-295.

Hall, G.W. and Huber, IhW., "System Description

and Performance Data for the USAF/CAL Variable

Stability T-33 Airplane," AFFDL-TR-70-71,

Wright-Patterson Air Force Base, OH, Aug. 1970.

Smith, R.E., "Effects of Control System

Dynamics on Fighter Apprmch and Lamling

Longitudinal Flying Qualifies - Volume I,"

AFFDL-TR-78-122, Wright-Patterson Air Fm_e

Base, OH, Mar., 1978.

Buethe, S.A., Govindaraj, ICS., and Knotts, L.H.,

"Description of NT-33A Configurations for Use in

Programming NASA Vertical Motion Simulator

(VMS)," Calspan Report No. 7205-5, Calspan

Corporation, Buffalo, NY, Mar., 1984.

DaneiG G., "Vertical Motion Simulator

Familiarization Guide," NASA TM 103923, May,

1993.

Tischler, M.B. and Canffman, M.G., 'T-xequen_-

Response Method of Rotorcrafi System

Identification: Flight Applications to BO-105

Coupled Rotor/Fuselage Dynamics," Journal of

the American Helicopter Society, Vol. 37, No. 3,

pp. 3-17, July, 1992.

10. Sinacori, J.B., "rhe Determination of Some

Requirements for a Helicopter Flight Research

Simulation Facility," NASA CR-152066, Sept.

1977.

11.

12.

13.

14.

15.

16.

17.

18.

Schroeder, J.A., "Helicopter Flight Simulation

Motion Requirements," Ph.D. Dissertation,

Stanford University, May, 1998.

Rolfe, J.M. and Staples, KJ., Flight Simulation,

Cambridge University Press, Cambridge, 1986.

Parrish, R.V., Dieudonne, J.E., Bowles, R.L., and

Martin, Jr., DJ., "Comdinatod Adaptive Washout

for Motion Simulators," Journal of Aircraft, Vol.

12, No. 1, Jan., 1975, pp. 44-50.

Martin, Jr., DJ., "A Digital Program for Motion

Washout on Langley's Six-Degree-of-_

Motion Simulator," NASA CR 145219, July,

1977.

Dieodonne, J.E., Parrish, R.V., and Bmdusch,

R.E., "An Actuator Extension Transformation for

a Motion Simulator and an Inverse Tt-amformafion

ApplyingNewton-Raphson'sMethod," NASA TN

D-7067,1972.

Cooper, G.E. and Harper, Jr., R.P., "The Usc of

Pilot Rating in the Evaluation of Aircraft

Handling Qualities," NASA TN D-5153, Apr.,

1969.

Myers, J.L., Fundamentals of Faperimental

Design, Allyn and Bacon, Inc., Boston, 1972, pp.

167-190.

Bray, R.S., "A Study of Vetticafl Motion

Requirements for Landing Simulation," Human

Factors, Vol. 15, No. 6, Dec., 1973, pp. 561-568.

11

American Institute of Aeronautics and Astronautics

485
q
A Method for the Flight Test Evaluation of PIO Susceptibility

Thomas R. Twisdale

412TW/TSFT / USAF Test Pilot School

Michael K. Nelson

USAF Test Pilot School

MNELSON@tps.edwards.af.mil

The handlingqualities test method taught at the USAF Test Pilot School is briefly described.

This method consists of three parts, or phases: Phase 1 is an evaluation of low bandwidth

handling qualities; Phase 2 is an evaluation of high bandwidth handling qualities; and Phase 3

is an evaluation of handling qualities during the operational tasks that make up the design

masston of the airplane. Phase 2 high bandwidth testing uses the Handling Qualities During

Tracking (HQDT) test technique, which when properly applied has proved remarkably

effective in exposing PIO susceptibility in airplanes of every,size and shape. For this reason

Phase 2 testing is often referred to as a handlingqualities "safety gate." If PIO is not

experienced during Phase 2 high bandwidth testing, it is unlikely that PIO will be experienced

during operational use. If high bandwidth handlingqualities are satisfactory, it is unlikely that

handling qualities will pose a significant safety of flight concern during operational use.

Introduction The three phase handling qualities

test and evaluation method described below has

been used at the AFFTC since 1972. When

used as described, it has proved remarkably

successful as a handling qualities test method

and as a means of "optimizing" the flight

control system to achieve improved handling

qualities. When used in a compromised fashion,

it has proved to be correspondingly less

successful. The second of the three phases,

which centers around high bandwidth Handling

Qualities During Tracking (HQDT) testing, has

proved especially successful in exposing PIO

susceptibility. Unhappily, this valuable tool has

often been misunderstood and misapplied, and

hence disparaged. Pilots who understand the

rationale for high bandwidth HQDT testing, and

who have been properly trained in the

specialized piloting technique, find it a very

effective handling qualities evaluation tool.

Discussion As all of flying qualities testing

should be, the three phase handling qualities test

method described below is grounded in the

model validation test method, which consists of

three steps:

. Predict the airplane response, based on

a model.

2. Test the prediction.

3. Validate or correct the model, based on

the test results.

The model validation test method is readily

recognizable as a form of the scientific method.

In Step 1, the handling qualities are predicted,

using available analytical criteria and piloted

simulators. We will not discuss Step I in this

paper. In Step 2, the airplane handling qualities

are tested using the three phase test method

described below. In Step 3, the handling

qualities model is validated. We will not

discuss Step 3 in this paper. The model

validation test method is the most effective, the

most efficient, and the safest way to conduct

testing. To fiarther emphasize test safety, the

handling qualities test method described below is

guided by the following procedural rule:

Employ a build-up approach, in which

testing progresses from the lowest to the

highest level of risk.

To ensure completeness, the handling qualities

test method described below is guided by the

following principle:

Handling qualities testing should explore the

entire spectrum of pilot-vehicle dynamics.

Betbre proceeding, we pause for two notes.

NASA Dryden PIO Workshop, 6-8 April 1999

487
First, we define handling qualities as the

dynamics, or characteristics, of the pilot plus

the airplane. Second, following the YF-22 PIe

incident, we at the Flight Test Center began to

refer to PIe as "pilot-in-the-loop" oscillation.

rather than "pilot-induced" oscillation. Pilots

must be in the loop for a PIe to occur, but

pilots do not induce these unwanted oscillations.

if anything, it is the airplane that induces them.

This is easily shown by noting that the same

pilot, flying two different airplanes in the same

manner may experience many PIes in the one

but never experience a PIe in the other. When

pilots understand that PIe is not their fault, they

are more likely to provide objective evaluations,

comments, and ratings.

The test method described below is composed of

three phases: a low pilot bandwidth phase, a

high pilot bandwidth phase, and an operational

phase. By "pilot bandwidth" we have in mind

both the range of frequencies and the amplitude

of control inputs generated by the pilot.

"Frequency content" would perhaps be a more

descriptive term, but "bandwidth" seems to be

more widely used. We will discuss each phase

of testing in turn.

Phase 1: Low Bandwidth Testing During

Phase 1 testing the pilot conducts an evaluation

of low bandwidth handling qualities at safe, upand-away flight conditions. By low bandwidth

handling qualities, we mean the handling

qualities characteristics that are associated with

relatively smooth (or low frequency), small

amplitude pilot inputs. We often refer to Phase

l testing as "warm-up," or "get acquainted," or

"familiarization" testing. Phase 1 low

bandwidth testing is designed to introduce the

pilot to the airplane under low risk conditions.

Phase 1 consists of relatively low bandwidth

piloting tasks, including open-loop tasks such as

pulse, doublet, and step inputs; semi-closed-loop

tasks such as low bandwidth pitch attitude and

bank angle captures, steady heading sideslips,

and so on; gentle maneuvering in the vicinity of

the test airspeed and altitude; and low

bandwidth, non-aggressive tracking.

You may object, correctly, that open-loop

maneuvers such as pulses, doublets, and steps

are not handling qualities test maneuvers at all,

because the pilot is not in the loop. We include

these maneuvers because they allow the pilot to

observe the dynamics, or characteristics, of the

airplane alone (even though experience shows

that an open-loop evaluation may be misleading

as an indicator of handling qualities).

Pilots must approach Phase 1 cautiously, even

though it is a low bandwidth evaluation.

Experience shows that airplanes with less than

desirable handling qualities may unexpectedly

and quickly draw a pilot into high bandwidth

control and PIe. For this reason, pilots must

focus on preserving low bandwidth, and be

prepared to relinquish control altogether (by

freezing or releasing the controls) to arrest an

unwanted response such as PIe.

When PIe, or other sufficiently undesirable

handling qualities are encountered during Phase

I low bandwidth testing, strong consideration

should be given to correcting these deficiencies

before testing progresses to Phase 2 high

bandwidth testing.

Phase 2: High Bandwidth Testing During

Phase 2 testing the pilot conducts an evaluation

of high bandwidth handling qualities. Most of

this testing is conducted at safe, up-and-away

flight conditions. By high bandwidth handling

qualities, we mean the handling qualities

characteristics that are associated with

aggressive, high frequency, small and large

amplitude pilot inputs. Phase 2 consists mainly

of HQDT testing. HQDT is perhaps the single

most important handling qualities test technique

at our disposal, especially when an evaluation of

PIe susceptibility is of interest. We often refer

to Phase 2 high bandwidth testing as a "safety

gate," because experience shows that when this

testing is executed correctly and PIe is not

exposed, the airplane may be considered PIefree with near certainty.

There are three principal components of I-IQDT

testing: the piloting technique, the test

NASA Dryden PIO Workshop, 6-8 April, 1999

488
maneuver, and the pilot evaluation.

The HQDT Piloting Technique The HQDT

piloting technique is a simple one. A small

precision aim point is selected on a target. This

aim point should not be larger than the pipper or

aiming index in the gunsight or head-up display.

The evaluation pilot's task is to track the

precision aim point as aggressively and as

assiduously as possible, always striving to

correct even the smallest of tracking errors as

quickly as possible. The effect of this simple

technique is to increase the bandwidth of the

pilot's control iriputs.

A systematic way to fully explore high

bandwidth handling qualities is to begin an

HQDT maneuver at low bandwidth (that is,

using small amplitude, low frequency inputs);

then increase the frequency range using small

amplitude inputs; then increase the input

amplitude while at high frequency. In practice,

you will find that this approach works well for

airplanes having satisfactory handling qualities,

but not as well for airplanes having less than

satisfactory handling qualities. The excessive

phase lag associated with degraded handling

qualities forces a pilot who is attempting to fly

with high bandwidth into a coupled pilot-plusairplane oscillation at a frequency below what

the pilot is capable of achieving. These lower

frequency coupled oscillations (which may or

may not be PIO) are ot_en a valuable indication

that the airplane handling qualities are not what

you would like them to be. In other words, the

inability to achieve high pilot bandwidth, despite

a vigorous attempt to dose, may itself be a

sign, in some cases, that the airplane handling

qualities are less than satisfactory.

Based on the description given in the preceding

two paragraphs, experienced pilots will

recognize that the HQDT piloting technique is

quite different from the low bandwidth

"operational" piloting technique used in normal,

everyday flying. In normal everyday flying,

experienced pilots do not resort to small

amplitude, high frequency inputs, and certainly

not to large amplitude, high frequency inputs.

Instead, they prefer small, smooth inputs deftly

applied in art effort to anticipate and correct

small errors before they grow into large ones.

Consider the operational "guns tracking" task, in

which an experienced pilot may initially lead the

target, then allow the gunsight pipper to drift

back to the target (or allow the target to drift up

to the pipper). Instead of aggressively

correcting tracking errors, relatively smooth,

measured corrections are applied with the goal

of "floating" the pipper toward the target. A

low bandwidth "operational" piloting technique

such as this will improve task performance

(especially when the handling qualities are less

than satisfactory), but it also hides, or masks,

the high bandwidth handling qualities of the

airplane. The purpose of the HQDT piloting

technique is to bring high bandwidth handling

qualities characteristics into the open, where

they can be evaluated.

Pilots who are unfamiliar with the purpose of

Phase 2 high bandwidth handling qualities

testing commonly raise several objections to the

specialized HQDT piloting technique. One is

that it is "unnatural," or "pilots don't fly that

way," or "HQDT might be okay tbr fighters,

but not for big airplanes because no one flys big

airplanes aggressively." A second objection is

that it results in degraded task performance. A

third objection is that "I can make any airplane

PIO" or "I can make any airplane look bad" by

using the HQDT piloting technique. A fourth

objection is that "we're only doing this to pacify

the engineers." The first objection is largely,

but not entirely true; the second objection is

true; and the third and fourth objections are

untrue. Let's look at each in turn, briefly.

The first objection, that the HQDT piloting

technique is "unnatural" in any airplane and is

inappropriate for large airplanes, is largely, but

not entirely true. Experience shows that the

HQDT piloting technique is not normally used

by pilots, but is an entirely natural response

when something happens to elevate a pilot's

level of excitement or anxiety above a certain

threshold. Also, the natural response of a

human pilot to high levels of excitement or

NASA Dryden PIO Workshop, 6-8 April 1999

489
anxietyisindependent of thesize of the

airplane. The space shuttle, the C-17, and the

B-2 are large airplanes, and each experienced

PIOs during testing. The second objection, that

the I,IQDT piloting technique results in degraded

task performance, is true. As a practical matter,

we observe from operational experience that

when excitement or anxiety precipitates a high

bandwidth response from a pilot, task

performance is degraded. The nature and level

of this degraded performance is of interest to us

in Phase 2 testing because it is one source of

incidents and accidents as well as degraded

mission pertormance. The third objection,

which is that "I can make any airplane PIO," or

"I can make any airplane look bad" by using the

HQDT piloting technique, is false. We show

the Test Pilot School students, first using a

simulator and then in flight, that a genuinely

Level 1 or Level 2 airplane cannot be made to

PIO. We show them that a Level i airplane

will feel crisp and responsive and follow their

commands closely even during high bandwidth

HQDT testing. They learn by experience that

the HQDT piloting technique will not make a

,ogp__dairplane look bad, but it will make a bad

airplane look bad. This, in a nutshell, is the

purpose of Phase 2 handling qualities evaluation:

to expose both the good and bad features of high

bandwidth handling qualities. The fourth

objection, which is that "we're only doing this

to pacify the engineers," is also false. Phase 2

testing, as all of handling qualities test{ng, is

conducted for pilots, not for engineers. It is

pilots, not engineers, who must fly the airplane,

perform the mission (sometimes under very

difficult circumstances that are conducive to

high pilot bandwidth), and return safely. It is

pilots, not engineers, who must live with the

consequences when the test community fails to

evaluate the full spectrum of handling qualities,

or tails to expose every deficiency, or fails to

correct deficiencies when warranted.

An interesting feature of the HQDT piloting

technique is that, in most cases, the evaluation

pilot is not allowed to use the rudder pedals.

This is referred to as "feet-on-the-floor"

tracking. At the Flight Test Center, experience

has taught us that much can be learned about

lateral-directional handling qualities when flying

feet-on-the-floor. Pilots are excellent aileron-torudder interconnects. When pilots are allowed

to use the rudder pedals, they can mask

handling qualities deficiencies that might

otherwise stand out prominently. However, the

HQDT piloting technique should not be thought

of as an exclusively feet-on-the-floor technique.

There are times when using the rudder pedals is

beneficial. For example, the pilot's description

of how the rudder pedals were used, together

with an analysis of the data, can be helpful in

correcting a deficiency.

In HQDT testing the evaluation pilot must not

be distracted by the measurement of task

performance, such as average tracking error, or

time within a given radius of the precision aim

point, and so on. Measuring task performance

encourages evaluation pilots to abandon or

compromise the I-IQDT piloting technique and

reduce their bandwidth. While reduced

bandwidth usually results in improved task

performance, it also compromises the evaluation

of high bandwidth handling qualities. When the

HQDT piloting technique is abandoned or

compromised, the average test pilot is quite

capable of producing good tracking results with

a pretty bad airplane. This tells us something

about the skills of the pilot, but it doesn't tell us

much about high bandwidth handling qualities,

which is what we are interested in during Phase

2 testing.

The HQDT piloting technique is not difficult to

learn, but it requires practice. The best place to

learn and practice this technique is in a flight

test simulator. Learning is easier and occurs

more rapidly when it is possible to estimate

power spectral density functions of the pilot's

control inputs immediately after a practice

maneuver.

We have noted the importance of large

amplitudes and high frequencies in high

bandwidth pilot inputs. By "high frequencies"

we do not mean that pilots should attempt to

track by generating high frequency sinewave

NASA Dryden PIO Workshop, 6-8 April, 1999

490
inputs. The high frequency component of high

bandwidth inputs comes from the sharpness, or

quickness of the pilots inputs. Sharp, quick,

control inputs are produced by reacting to

tracking errors as rapidly as possible.

We must emphasize the importance of an honest

and vigorous effort to use the specialized, high

bandwidth, HQDT piloting technique.

Otherwise, high bandwidth handling qualities

(which are usually the worst handling qualities)

will not be fully evaluated during the test

program. Instead, these handling qualities will

be evaluated in the field, during operational use

by line pilots rather than test pilots.

We conclude our brief description of the

specialized HQDT piloting technique by

remarking again that this technique, which lies

at the heart of Phase 2 high bandwidth testing,

is often compromised by pilots and engineers

who regard it as unnatural and artificially

contrived. In fact, however, this technique is

entirely natural under certain circumstances.

You need only examine time histories of pilot

control inputs during a Pie to see that this is so.

HQDT Test Manuevers The heart of high

bandwidth handling qualities testing lies in the

specialized HQDT piloting technique. Any

maneuver that requires the evaluation pilot to

use the specialized, high bandwidth, HQDT

piloting technique is likely to be a suitable

HQDT test maneuver. For this reason there is

no exclusive catalog of HQDT maneuvers.

Maneuvers that have worked well in the past

include constant load factor (or angle of attack)

air-to-air tracking maneuvers, wind-up turn

tracking maneuvers, tracking while closing on

the target, tracking in the power approach

configuration (with and without closure), air-toground tracking, refueling boom tracking, and

formation flying. Other maneuvers, perhaps

better suited to a particular airplane, may be

invented as the need arises.

Formation maneuvers and refueling boom

tracking maneuvers should not be flown so close

to the lead airplane or to the refueling boom that

the evaluation pilots feel that their safety is

compromised by the high bandwidth HQDT

piloting technique of aggressive, assiduous

tracking.

With a single exception, a fixed pipper or

aiming index is used during HQDT testing.

When a moving pipper or aiming index is used

(as in the case of a computing gunsight), the

pipper (or gunsight) dynamics become a part of

the evaluation. Our initial goal is to evaluate

the dynamics of the pilot plus the airplane,

rather than the pilot plus the airplane plus the

gunsight. Hence a fixed pipper is nearly always

used. The exception arises later, when it might

prove desirable to evaluate the effect of the

computing gunsight dynamics on handling

qualities. Used in this way, HQDT can be an

important tool for fine-tuning the gunsight

component of the pilot-vehicle dynamics.

The depression angle of the pipper or aiming

index is usually dictated by the airplane and the

test maneuver. The depression angle may be set

to minimize pendulum effect; or set to the angle

that would be computed by the gunsight for a

given load factor (in air-to-air tracking) or for a

given dive angle (in air-to-ground tracking); or

set to aid in avoiding the target airplane jetwake.

The test airplane must not be retrimmed during

the test maneuver. Trimming detracts from the

pilot's concentration on high bandwidth tracking

and renders invalid a frequency response

analysis of the test data (unless the trim imputs

are recorded and made available for analysis).

Pilot Evaluation Pilot evaluation is the third

component of Phase 2 HQDT testing. In HQDT

testing, pilot comments are the most important

part of the pilot evaluation, supported by a PIO

rating. Careful and complete pilot comments

from HQDT testing are the key to helping

designers and flight test engineers understand

the high bandwidth handling qualities of the

airplane. Cooper-Harper ratings are not

assigned following an HQDT evaluation because

task performance (such as tracking error) is not

measured during HQDT testing. Hence, it is

NASA Dryden PIe Workshop, 6-8 April 1999

491
notpossibleto assignalegitimateCooperHarperatingbased on an HQDT evaluation.

Phase 3: Operational Testing During Phase 3

testing the pilot conducts an operational

evaluation of the airplane handling qualities.

The purpose of Phase 3 testing is to determine

whether the handling qualities are suitable for

pertbrming the various tasks that make up the

design mission. Depending on the airplane.

these tasks may include take-off, landing, aerial

refueling, formation flight, and air-to-air and

air-to-ground weapons delivery. Phase 3

operational testing must often be conducted in

the presence of aggravating factors such as

atmospheric turbulence, darkness, proximity to

the ground, and so on. The risks associated

with these factors must be explored in a buildup fashion. Cooper-Harper ratings are

appropriate during Phase 3 operational testing.

Conclusion The overarching objective of the

three phases of testing we have briefly described

is to completely evaluate the full spectrum of

airplane handling qualities. When we fail to

achieve this objective, operational pilots become

test pilots by default, but without the necessary

preparation and safeguards we bring to bear in a

properly conducted flight test program. For this

reason, the entire range of handling qualities

must be explored by test pilots during flight

_esring, rather than by operational pilots during

operational use of the airplane.

At present, Phase 2 high bandwidth testing using

HQDT test techniques is perhaps the most

important tool we have for evaluating high

bandwidth handling qualities characteristics,

particularly PIO susceptibility. HQDT testing is

often resisted or disparaged because its purpose

and rationale are not understood, or because it

has been used incorrectly by pilots who were

not properly trained in the specialized HQDT

piloting technique. When used properly, I-IQDT

has proved to be uniquely successful. Properly

conducted Phase 2 high bandwidth HQDT

testing serves as a handling qualities "safety

gate." If high bandwidth handling qualities

prove to be satist_ctory, it is unlikely that

handling qualities will pose a significant safety

of flight concern during operational use of the

airplane. It"PIO is not experienced during

HQDT testing, it is unlikely that it will occur

during operational use.

NASA Dryden PIO Workshop, 6-8 April, 1999

492
Pilot Opinion Ratings and PIO

Michael K. Nelson

USAF Test Pilot School

MNELSON@tps.edwards.af.mil

Thomas R. Twisdale

412TW/TSFT / USAF Test Pilot School

Two simple measures for dramaticallyimproving the assessment of PlO susceptibility are

presented, together with supporting arguments. These measures are first, to welcome, rather

than suppress, the exposure of PIO susceptibility; and second, to assign a Cooper-Harper

rating of 10 to every PIO, whether fully developed or incipient. A Cooper-Harper rating of

10 is a declaration that the airplane is uncontrollable during a PIO. It is argued that such a

declaration is reasonable because pilots must necessarily relinquish control, if only

temporarily, in order to arrest a PIO.

Discussion For more than 25 years, it has been

possible to obtain reliable flight test assessments

of PIO susceptibility using available test

methods and rating scales. Yet reliable

assessments are not the rule. We believe they

could be made the rule by adopting two simple

measures:

1. Welcome, rather than suppress, the

exposure of PlO susceptibility.

. Assign a Cooper-Harper rating of lO to

every PIO, whether fully developed or

incipient.

We grant that adopting these two measures

would require overturning long standing, deeply

ingrained practice. But our experience suggests

that traditional practice is misguided and

counter-productive. We will discuss each of

these proposed measures in turn.

Welcome the Exposure of PIO Susceptibility

PlO is not welcome during a flight test

program. Consequently, pilots are under subtle

but significant informal pressures to ignore,

overlook, play down, or explain away

occurrences of PlO. The reasons for these

pressures are well known: a strong desire to

maintain a success-oriented test schedule and

budget; the fear of Congressional scrutiny; the

fear that Congress will cancel a needed airplane,

and so on. Because of these pressures an

encounter with PIO can, in our experience, lead

to a variety of pilot assessments. If the airplane

is damaged or lost, the pilot would likely agree

that a PIO occurred and a Cooper-Harper rating

of 10 might be assigned (although in flight

testing such a rating is uncommon). If the

airplane is not damaged or lost, the pilot might

not mention the PIO at all. Or the pilot might

initially acknowledge that a PIO occurred, but

later deny it. Or the pilot might acknowledge

the PIO, but blame it on himself. (How many

times have experienced handling qualities testers

heard a pilot say: "I screwed up. If I hadn't

.... I wouldn't have gotten into a PIO.")

Occasionally, a pilot will acknowledge the PIO

and suggest that the airplane needs to be fixed,

but the pilot who offers this assessment often

suffers for his honesty.

We believe that the discovery of handling

qualities deficiencies of every kind, including

PIO, should be welcomed. The purpose of an

acquisition program is to provide the operational

users with an airplane that is suitable for

performing the various tasks that make up the

design mission. Line pilots rely on the test

community to evaluate handling qualities

thoroughly and objectively. They rely on the

acquisition community to correct those

deficiencies that warrant correcting (those that

render the airplane unsafe or less than suitable).

But these deficiencies cannot be corrected if they

have not been found, or have been ignored or

played down. Handling qualities deficiencies

should be discovered by test pilots during the

test program, not by line pilots during

operational use. Test pilots should be given to

NASA Dryden PIO Workshop, 6-8 April 1999

493
understandthatit ispartof their job to discover

strengths and deficiencies, and they should be

lauded when they do. The discovery of an

important deficiency should be regarded as an

opportunity to provide a better finished product.

We should note in passing that following the

YF-22 PIO incident, we at the Flight Test

Center began to refer to PIO as "pilot-in-theloop" oscillation, rather than "pilot-induced"

oscillation. Pilots must be in the loop for a PIO

to occur, but pilots do not induce these

unwanted oscillations. If anything, it is the

airplane that induces them. This is easily shown

by noting that the same pilot, flying two

different airplanes in the same manner may

experience many PIOs in the one but never

experience a PIO in the other. When pilots

understand that PIO is not their fault, they are

more likely to report occurrences of PIO and

provide objective evaluations, comments, and

ratings.

At present, PIO susceptibility is not always

adequately explored and reported because test

pilots and engineers recognize that PIOs are not

welcome news. Perhaps the most effective way

to immediately improve the assessment of PIO

susceptibility is to welcome encounters with PIO

during flight testing.

Assign Cooper-Harper Ratings of 10 to Every

PIO We believe every PIO, whether fuUy

developed or incipient, should be assigned a

Cooper-Harper rating of 10. This is equivalent

to saying that every PIO, whether fully

developed or incipient, represents at least a

temporary loss of control. We define fully

developed and incipient PIOs in the following

way. A fully developed PIO is one in which

several cycles of the oscillation occur, even

though the oscillation may not reach a visibly

steady state. An incipient PIO is one which the

pilot is able to recognize and quickly arrest,

perhaps within a cycle or less.

Some in the handling qualities flight test

community would agree that a fully developed

PIO india:ares a loss of control, and therefore

warrants a Cooper-Harper rating of 10. But

many would disagree, contending that when the

pilot is able to arrest a fully developed PIO and

continue with the task, control has not been lost,

at least not in a long term, or global sense.

They would further contend that a CooperHarper rating of 10 is warranted only when the

PIO results in a stall, departure, collision with

another airplane or the ground, or complete

abandonment of the task. Few in the test

community would agree that an incipient PIO

warrants a Cooper-Harper rating of 10. If it

can be shown that both fully developed and

incipient PIOs represent a loss of control, then

perhaps we can agree that every PIO should be

assigned a Cooper-Harper rating of 10. We will

turn our attention first to fully deveoped PIO,

then to incipient PIO.

F_llY Developed PIO Let us first explore the

question of whether a fully developed PIO

represents a loss of control. We begin by

asking how a pilot arrests a fully developed

PIO. One of three methods is usually

employed: the pilot either freezes the controls,

or releases the controls, or significantly reduces

bandwidth (or the aggressiveness of control).

When a pilot freezes or releases the controls, he

has clearly relinquished control of the airplane

for a time sufficient to arrest the PIO. Does it

not follow that the pilot has also abandoned the

task during the time required to arrest the PIO?

While the controls are frozen or released, the

pilot cannot be tracking the target, or controlling

the flare, or whatever. If this is the case, we

may ask why the pilot has abandoned the task if

he still has control over the airplane. Isn't the

answer that the airplane was uncontrollable

during the PIO? When a pilot significantly

reduces bandwidth to arrest a PIO, we would

suggest that he has, in effect, transitioned from

the primary task (tracking, landing, refueling,

and so on) to the suddenly more important task

of regaining control. We would even suggest

that significantly reducing bandwidth is really

another form of temporarily freezing the

controls.

Implicit in our discussion is the understanding

NASA Dryden PIO Workshop, 6-8 April, 1999

494
that when a pilot temporarily relinquishes

control to arrest a P[O, he does so as a matter

of necessity rather than choice. If it is

necessary, for the pilot to relinquish control in

order to arrest a PlO and reestablish control,

aren't we acknowledging that the airplane was

temporarily uncontrollable? If the airplane was

controllable, why did the pilot find it necessary

to relinquish control?

Nevertheless, the objection will be raised that if

a task is performed one hundred times and PIO

is encountered only once, it would be silly to

claim that the airplane is uncontrollable. We

believe the proper rejoinder to this objection is a

reminder that Cooper-Harper ratings are

assigned to individual evaluations, or trials. If a

PlO was experienced only once in one hundred

evaluations of the same task in the same

configuration at the same flight conditions, we

would argue that the pilot lost control only once

in one hundred evaluations, and that the airplane

proved to be uncontrollable only once in one

hundred evaluations, so that a rating of 10 was

warranted only once in one hundred evaluations.

This one data point out of a hundred is an

important one that should not be swept under

the rug or played down. If it can happen to a

test pilot once in a hundred times, how often is

it likely to happen to less experienced and

possibly less skilled line pilots?

Incipient PIO Now let us turn our attention to

the question of whether an incipient PIO

represents a loss of control. In Figure 1 we

present a sketch comparing time histories of

pitch rate response and stick force during two

events of interest. In one event, represented by

dashed line time histories, we see a fully

developed PlO. In the second event,

represented by solid line time histories, we see

an incipient PIO. Both PIOs were precipitated

by identical circumstances. At the first arrow,

nose down pitch rate begins to develop and the

pilot counters by nudging the stick aft, but

without apparent effect (perhaps because of

excessive phase lag), so that nose down pitch

rate continues to increase. The pilot continues

to smoothly increase countering stick force until,

suddenly, at the second arrow the airplane

begins to pitch up rapidly. In an attempt to

arrest this rapid and unsettling reversal of

motion the pilot takes action. In the P[O

represented by the dashed line time histories, the

pilot makes a moderately large and rapid control

inputin the opposite direction, which aggravates

the airplane response and causes the pilot to

transition from low to high bandwidth control.

A fully developed P[O ensues. In the PIO

represented by the solid line time histories, the

pilot adopts a different course of action.

Recognizing that a PIO is about to begin, the

pilot makes a small corrective input to arrest the

unwanted motion and then relinquishes control

by freezing the stick. After a short interval

(perhaps a second or two, perhaps only a

fraction of a second), the pilot gets back into the

loop and resumes flying the airplane. Note that

there is no visible evidence of PIO or PIO

susceptibility in the solid line time histories of

this incipient PIO. Only the pilot is aware that

he intentionally relinquished control in order to

avoid the PIO he sensed was about to ensnare

him. When flying an airplane that is PlO

susceptible, it is not uncommon for pilots to

repeatedly relinquish control to forestall PIO.

nose-tip ]

J I ! •

I

nolle- |

down I ',._.s/ "._ _/

Figure I A comparison of two eventS.

What distinguishes the events recorded in the

dashed line time histories from those recorded in

the solid line time histories? In the dashed llne

time histories, the pilot remained in the loop and

allowed himself to be drawn from low

bandwidth control into high bandwidth control,

resulting in a fully developed PIO. In the solid

NASA Dryden PIO Workshop, 6-8 April 1999

495
line time histories the pilot made a small

corrective input and then temporarily

relinquished control of the airplane until the

unsettling motion subsided, thereby avoiding any

visible evidence of PIe or PIe susceptibility.

But we see in this second case that the pilot did

embark on a PIO, before quickly arresting it by

temporarily relinquishing control. In other

words, a PIe was encountered in both cases: in

the one, the PIe became fully developed,

whereas in the other the PIe was incipient. In

both cases, we believe the airplane should be

described as PIe susceptible.

Most pilots and engineers would argue that the

event recorded by the solid line time histories in

Fibre [ is simply an example of pilot

compensation, and indeed we acknowledge that

this is so. By temporarily relinquishing control

(a tbrm of compensation), the pilot succeeded in

arresting the PIe at the incipient stage, betbre it

could become fully developed. As every

experienced pilot knows, when an airplane has

poor handling qualities, temporarily

relinquishing control can be a very effective

form of pilot compensation. Skilled.

experienced pilots know when to exercise

control and when to leave well enough alone.

When poor handling qualities are in evidence, it

has been observed that the best pilots are those

who exercise the most forethought and the least

control. Unfortunately, this form of

compensation may hide serious deficiences from

everyone but the pilot, who may choose not to

mention them. Our concern is that, by

regarding the temporary relinquishing of control

as compensation, the pilot is hiding the fact that

an airplane is PIO susceptible. We believe that

when control must, of necessity, be temporarily

relinquished to arrest or forestall PIe, whether

incipient or fully developed, the airplane must

be regarded as temporarily uncontrollable. To

regard it otherwise is to risk assessing the PIe

susceptibility of pilots rather than airplanes.

Conclusion For more than 25 years, it has

been possible to obtain reliable flight test

assessments of PIO susceptibility using available

test methods and rating scales. However, many

pilots and engineers have deduced from flight

test practices that PlO encounters are

unwelcome. Available test methods and rating

scales are not always used, or are used in a

compromising manner, rendering them less

effective: and subtle pressures may be brought

to bear on pilots, encouraging them to ignore,

overlook, play down, or explain away PIe

encounters. We are presently quite capable of

thoroughly and accurately assessing PIe

susceptibility, but we believe that such

assessments will not become routine until two

simple measures are adopted: first, welcome

the exposure of PIe; and second, assign a

Cooper-Harper rating of 10 to every encounter

with a PIe. whether fully developed or

incipient.

To some, it will seem Procrustean to insist that

every occurrence of PIO be assigned a CooperHarper rating of 10. After all, this is a

declaration that the airplane is uncontrollable,

which is a harsh word. Nevertheless, the

strategy for arresting a PIe is to temporarily

relinquish control, which leads us to the

que._tion: if an airplane is controllable, why

should it ever be necessary to relinquish

control? When control is given up of necessity,

doesn't this mean that the airplane could not be

controlled, and is therefore uncontrollable, even

if only temporarily? Although the strategy of

temporarily relinquishing short term control in

order to preserve long term control may

legitimately be described as pilot compensation,

doing so serves to camouflage PIe

susceptibility. The pilot may recognize what he

is doing, but he is unlikely to mention it to

anyone else.

NASA Dryden PIe Workshop, 6-8 April, 1999

496
REPORT DOCUMENTATION PAGE I Form OMB No. Approved 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

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4. TITLE AND SUBTITLE

Pilot-Induced Oscillation Research:

Status at the End of the Century

6. AUTHOR(S)

Compiled by Mary F. Shafer and Paul Steinmetz

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

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13. ABSTRACT (Maximum 200 words)

The workshop "Pilot-Induced Oscillation Research: The Status at the End of the Century," was held at NASA

Dryden Flight Research Center on 6-8 April 1999. The presentations at this conference addressed the most

current information available, addressing regulatory issues, flight test, safety, modeling, prediction, simulation,

mitigation or prevention, and areas that require further research. All presentations were approved for publication

as unclassified documents with no limits on their distribution. This proceedings includes the viewgraphs (some

with author's notes) used for thirty presentations that were actually given and two presentations that were not

given because of time limitations. Four technical papers on this subject are also included.

i

14. SUBJECTTERMS

Flight control, Flight safety, Pilot-induced oscillation, Simulation of flight test

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