NAS A/CP-2001-210389/VOL 1

Pilot-Induced Oscillation Research:

Status at the End of the Century

Compiled by Mar 3, E Shafer and Paul Steinmetz

NASA Do, den Flight Research Center

Edwards, California

April 2001
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NASA/CP-2001-210389/VOL 1

Pilot-Induced Oscillation Research:

Status at the End.of the. Century

Compiled by Mary E Shafer and Paul Steinmetz

NASA Dryden Flight Research Center

Edwards, Califonua

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273

April 2001
NOTICE

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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

FOREWORD ..................................................................... iii

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-lnduced 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 (3. 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 (3. 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 ftir 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-Induced 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 I

Modeling tile Human Pilot i,I Single-Axis

Linear & Nonlinear Tracking Tasks

Yasser Zeyada, and Ronald A. ltcss

Dept. of Mechanical and Aeronautical l:.ngineeritlg

tlniversity of C:difomia

I)avis, CA 9561_

Outline

Inhoduction

Analytical Approach

Structural Model

Linear Analysis (Program PVD)

- Nonliqear Analysis (Program PVDNL)

huproved Version of PVDN_ with Graphical IJ_r Interface

Analyzing HAVE LIMITS data

Design Example - I .ongitudinal Flight Control System For HARV

Self-Report Card on "Criteria for Criteria"

Conclusions
Introduction

Mt_i_ allan

- "'Rcsuarch to devulop design ass¢._snlenl crileria ill]d ;ul:llysis h_ols should

focus on Calegory II and III PIOs....This research should ¢t_mbinc

experiments with the dcvclopmcnl ol'clli:clivc malhemulical analysis

mclhods capable of rationalizing and emulating the cxperinncntal rCSLIIIs"

_) R_OIIIIIIClKIIILiUII (J-_ ,_t'l_lllt/ll ,_"<,_/L'IV4#Id IJlhJl I "olllrr*l, Repo01 *_1"Ihc

COlllUlllli¢¢ 011 Ih¢ EITec.ls or Aorcral'l-I_h_! COUl)lil_L_ n,i Flit'lit galL'ly, NRC,

1991

Approach

I-xlend linear, closed-loop, HQWPIO p0"¢diclion technique to vehicles

with sigilil]cant no=dincaritics, e.g., acl.ualor l_tlc saturation

Assess technklue using HAVE LIMITS flight test data

Analytical Approach

Principal Assertions

Aircrall handling qualities, including PIO events arc fundamentally closcdIt)op phcIlO, IlICfla

A tmil_'ing Ihcory for h_mdling qualities and PIO, should, thercrorc, adopl a

closcd-looo perspective

A closed-loop perspective, of necessity, requires ;Lmodel or the human pilol

4
Analylical Approach

{.- -It.

Struclural Model ol'l-luman I'ilL_t

I'll.&_}

r- ................. ]

@-u L:_"-

Analytical Approach

Structural Model of Human Pilot

"Regressive Mode" - Assumed to Occur in Fully-Developed PIO

I'll.o]*

,,,1.J_,v,; i

"'1'-- "=_] ..._._. ,--, ',F,---_-7-_."_"_

1.-

vl._u.I /e¢, lt_ac _t

5
Analytical Approach

Applying Structurul Model to l,incar Vehicles

Methodology developed in

- I less, It. A., "Unil_'ing Theory for Aircraft 11_mdling Qualities and

Adverse Aircraft-Pilot Coupling," Journal of Gztidtmce. ('onlrot, atrd

Dynamics, Vol. 20, No. 6, 1997

Interactive MATLAB-based computer program developed as

- Zeyatda. Y., and Hess, 17,. A. "PVD Pilot Vehicle _D_3numics, An

Interactive Computer Program for Modeling the Human Pilot in SingleAxis I.inear ]'racking Tasks, Dept. of Mechanical and Aeronautical

Engineering, UC Davis, 1998.

Analytical Approach

The H,'mdling Qualities Sensitivity Function (14QSF)

Given model of vehicle dynamics, PVD allows creation of a Structural Model

of the pilot

The tlQSF is defined by I Ur,¢'C t, after normalized by gain K_ in model

Using NT-33A and TIFS flight test data, bounds on I UM/C I obtained which

could delineate handling qualities levels

li,

II(2*;F J e_YI ¢

-- x

j •"

' /, L-_i-_.:---

I . "_ • LrI'rl I

i ._ t ..... _ .... I ......

2 4 6 fl I0

6
Analytical Approach

The Power Spectral Density of U M (_,..,,_(t0))

Given model of vehicle dynamics. PVD allows creation of a Structural Model

of lhe pilot

The power spectral density of UM, alter normalized by gain K=z in m(_dcl, is

obtniucd

Using N'I'-33A and TIFS flight test data, bounds on (1),,, u,,,0t_) obtained which

cot,ld delincatc PIOR "lcvels"

4

Oo (_)

=u.

3

"X PlLll_ ,_ 4

x

I

\

2 _ PIOR < J

I r I I 1

1 2 3 4

Analytical Approach

Example - A LAHOS Config. with 0.2 s time delay added

Handling Qualities Level Pilot-Induced Oscillation "Lever"

J

l/ h-J

D!....... - ............
Analytical Approach

Applying Structural Model to Nonlinear Vehicles

("Nuisance" Nonlinearities)

Methodology developed in

- Hess, R. A., a,ld Stout, P. W., "Assessing Aircraft Susceptibility to

Nonlinear Aircraft-Pilot Coupling/Pilot-Induced Osci/lations, Joltrnctl o[

Guidanc'e, Control and Dyntvnics, Nov.-Dec. 1998, pp. 957-965)

Interactive MATLAB/Simulink-based computer program developed as

Zeyada, Y., and tless, R. A., "PVDNt Pilot/Vehicle _D_ynamicsv4,,,,Li.... An

Interactive Computer Program for Modeling the Human Pilot in SinglcAxis Linear and Nonlinear Tracking Tasks, Dept. of Mechanical and

Aeronautical Engineering, UC Davis, 1998.

Analytical Approach

No fundamental changes in theoretical approach .... normalized HQSF

and tl),,,(_o) still used, but obtained from ,ionli,lear Simulink simulation

HqbSF now obtained as

u (z)e -._,.,o

H(?SFclt)e "¢"_d

Io t,...,

....(e))now obtained as

i - 1,2....,50

¢_"'(_) 1,_ • 4 I
Analytical Approach

Example - A LAHOS Config, with amplitude and rate-limited elevator actuator

Handling Qualities Level

i ......

Pilot-Induced Oscillation "Level"

' i

Improved Version of PVDNL with GUI

_.____1'_1,._1 *" fl

..... __; - '_.; ....

9
hnproved Version o1"PVI)N;" with GUI

'1

o!!....

L m_ a

Improved Version of PVDM. with GUI

-- V,mll_

_ NLaV-J[

.OL'!_ I t )

mO_

-,_ _mpr"

i J 4

10
HAVE LIMITS Flight Tests

USAF-Sponsored flight tests using (for the last time) the NT-33A variable

stability aircraft

Goal: Evaluation of effects of actuator rate limiting on longitudinal handling

qualities and PIO

Three ¢onl'_guralions evaluated:

- 2D (stable unaugmentedairframe

- 2P (essentially 2D with stick filter)

- 2DU (unstable unaugmented airframe, similar to 2D when augmented) _'

Two HUD pitch-attitude commands utilized

- sum ofsinusoids

- discrete, step-like _¢/

HAVE LIMITS Flight Tests

(Pilot 3)

Cooper-Harper Rating

10

987 i 2DU2D2P

+ !

2 o___2/

1 6

t I I I l l + _..I

10 20 30 40 50 60 157

actuator rate limit

(dee/s)

Pilot-Induced Oscillation Rating

6

5

4

3

2

I

I

10

i i

I I J J _,,v..j

21) 30 40 50 60 157

actuator rate limit

(deg/s)

11
Analyzing HAVE LIMITS data

Configuration 2DU with rate limit = 157 deg/s

itandling Qualities Level Pilot-Induced Oscillation "l.evel"

6;

t /lOaF s k_t I

I

f i

°*t_-- ;"- ; 7 i ; _ _ 7 i %

I >

Flilhl 1_ PIOII = 7_

s t2/)) ¢no_.. 4

Analyzing HAVE LIMITS data

Configuration 2DU with rate limit = 60 deg/s

(pilot/vehicle system unstable @ 40 deg/s)

Handling Qualities Level

_ _, miihc Tom HQF - IO,

Pilot-Induced Oscillation "Level"

Fll_ITcsl PK_ - _ t,mll,= 1

i!

12
Analyzing HAVE LIMITS data

Configuration 2DU with rate limit = 53 dcg/s

(minimum rate limit tot pilot/vehicle stability)

Rate-tracking Structural Model Predicted (fully-developed) PIO

I ................ 1

hi, A t

r_. ]----_' I I er_----[ .......

,-'__, - __,,b _' " ;=-'_ _-IT.....

............................. I

lo_--_f,--i lbjq '

e_

_o = 5.4 rad/s ---J'

2 3 4 5 e

-- w._l¢

Aj' ,j

Design Example

Longitudinal Control of HARV

Control structure

Reduced-order model

- only rigid-body vehicle dynamics considered - (dynamics of two actuators

ignored)

- simpt¢ two-state reduced-order model results (short-period vehicle model

used)

13
Nonlinear Pilot/Vehicle Analysis

Actuator rate and amplitude limiting must be considered in final handling

qualities evaluation

Pilot/vehicle system

Pitch command

÷ V----q +, r--z--_-_

0c_ 0

;I+,-, _ i

:!i¢ll L/

Nonlinear Pilot/Vehicle Analysis

Initial predicted handling qualities and PIO levels

using Structural Pilot Model and program PV D_

Flight Cond: Math No. = 0.3, All. = 26.000 ft

full :t:20% perturbations on vehicle A, and Br matrix elements

Iri_ml,_ ? rad.'l

_ _ %r,

i _l it i._ :l _s i +s s

llrqwnc# rldl/i

14
Nonlinear Pilot/Vehicle Analysis

Predicted handling qualities and PIO levels

after addition of anti-windup logic in GQFr(S)

,[. tt-vttJ

o ; 2 2 * _ i 7 Ib It

Fi'_mnfy mafit

I,Iro_ct, lltl,, pe.w_

Self-Report Card on

Criteria for Criteria

Definitions taken from NRC PIO report

• Validity: Implies that a criterion embodies properties and characteristics that

define the environment of interest...criterion must rel_lte to closed-loop, highgain, aggressive, urgent and precise pilot-control behavior

Grade = 7.5/10

• Selectivity: Demands that criterion differentiate sharply betweesl "good" and

"bad" systems.., in context of PlO prediction, must distinguish between

configurations that may be susceptible to severe PlOs from those that are not

Grade = 7II 0

• Ready Applicability: requires that criterions be easily and conveniently

_a/,plic,/

Grade = 6.5/10 (Original PVDNL )

=7.5/10 (PVDr, rt" with GUI)

15
Conclusions

• Unifying theory for handling qualities and PIO can be ofli:rcd fi)r both linear

and nonlinear (nuisance nonlinearity) systems

Structural Pilot model, implemented in a computer-aided design program

provided predictions or" handling qualities levels and PIOR levels which

compared well with those from ltAVE LIMITS flight tests

• Methodology could be said to receive passing grade in "Criteria for Criteria"

16
Bandwidth Criteria for

Category I and II PIOs

David G. Mitchell

Hoh Aeronautics, Inc.

David H. Klyde

Systems Technology, Inc.

Pilot Induced Oscillation Research Workshop

NASA Dryden Flight Research Center

I'1_"'''a|_ '_|--- 6 April1999

Background

• Phase II SBIR from Air Force Research Labs

- Development of Methods & Devices to Predict & Prevent PIO

- Contract monitor is Tom Cord

- In process of writing final report

• Goals:

- Gather data (Lockheed Martin, Northrop Grumman,

McDonnell Douglas subcontractors)

- Analyze all available PIO data

- Develop criteda for prevention by design

- Develop test methods for detection in flight test

- Develop devices for real-time monitoring and detection

I?
Outline

• Pitch criteria based on airplane Bandwidth for

- Handling qualities

- PIO

• Apply research, experimental, operational data

• Compare Smith-Geddes, Gibson, NeaI-Smith criteria

• Bandwidth criteria for Category II PIO

• Control/response sensitivity and PIO

• Extension to roll axis

• Recommendations

Analytical Criteria

• Category I PIOs (linear):

- Many criteria exist

- Bandwidth-based criteria show most promise

• AIAA-98-4335 show them to be effective

• Amenable to initial design through flight test

• Category II PIOs (rate limiting):

- Only a handful of criteria

- Most are complex to apply

• Require dosed-loop analysis

• Applicable to analytical models only, not in flight

• Must make assumptions about pilot, fTequency, or amplitude

- Recent work on Bandwidth criteria shows promise

18
Handling Qualities Criteria

° Criteria developed for draft MIL standard (AFWALTR-82-3081, 1982)

- Requirements more stringent than "classical"(CAP) criteria

- Almost didn't make it into MIL-STD-1797 (1987)

• Primary short-term response criteria in rotorcraft

handling-qualities standard ADS-33D-PRF

• For airplanes, adopted revised version of Gibson's

requirements on dropback/overshoot

- Relaxed Bandwidthlimits(WL-TR-94-3162)

- USAF TPS projectfounddropbackuntestableinflight

(AFFTC-TR-95-78)

- Dropbacksecondaryinimportanceto pitchrate overshoot

- Current criteriause frequency-domainmeasureof overshoot

Process for Obtaining Bandwidth

Information from Flight

F_quency sv_ep .... Bade plot...

2o - : i =*=.m

• , J ,

19
Attitude Bandwidth Parameters

'_S_(wB%Vga& _, WSWphs,¢ ) !

(_)

-270

_o

Fmuency - k_ ecekt (rsdh_¢)

Pitch Rate Overshoot

t#.q I [q!tlt \

(dB) __ --_[_

0.01 0.1 1 1

2O
Nonlinearities Can Cause Data

Quality to Degrade

• Example data from inflight frequency sweep

• Coherence drops as a

result of rate limiting

_ p2is a measureof linear

correlationbetweeninput

and output

• Input power high

• Frequency response

looks reasonable

• Examined in AIAA-99-

0639 (Reno)

2t:

I_e (din 0

I

I)8

•0_-o,,,,,_..... ", , ,':']'_

.__._ ,.?..']...

21;6 * " *°+e'**'  l

P o4

02 _**  =.,.

t 10

Bandwidth Criteria for Handling

Qualities (Fighters -- Landing)

o.4[

Level 3

o,I

Pr.me

C_ay. 0.2

0, ! Level 2 _1 1

°J

0 ! l 3 4

HAL+_"_ ----'"'-'

S 0

21
Bandwidth Criteria for PIO

(Fighters-- Landing)

Prom

w

o3  ̧

Level 3

"Severe" PIO

O2

t'_) _._..,,_""_evel 2 "Mild" PIO if Flight Path Bandwidth

=BWT< 0.7red/sec J

I Level 21 Level 2 J

o.1! "Mild" I No PIO / Level 1

PIO II ISobbk' I NO PIO

if&G(q)ll if AG(q) I [Pitch Bobble if Pitch Rate

o >12dBII >9dB|, i OvershootRatio&G(q)>gdB]

0 1 2 3 4 S 8

P_ Amwdl _. mBW8 {tad/No)

Criteria Applied to Research Data

Successful on 188 of 207 (91%) [78 of 91 PIOs (86%)]

o T 2 3 4 5 II

22
Gibson Criteria (Research Data)

166 of 207 cases (80%) [66 of 91 PIOs (73%)]

• ] a Tendency I (Some cases fail aUude

200 %. _ ----_ j gain r'_qumr_nt)

m

I_//" L3 (PIO)

m=/

Average 1¢'° _ _"

_r_ z .]m_ ;mmm| L2 (P,O) l Note: , = (avg phase rate)Y720

( )t_ = =o

,o °8° o

i U(oooo.,_=o o o

L_O o c2 °_omm_O _ eoo ==d_aoo_¢,#O o o

Frequency at 180 deg lag (Hz)

NeaI-Smith Criteria (Research Data)

158 of 207 cases (76%) [75 of 91 PIOs (82%)]

re6onsnce (dB) •

o

0"

p •

oo • •

• _m o_mm m

¢p °o •

D )IB.,,m'_ • o am Oo • •

_1_ o qe, •

• _ . ,LWal • O0 ir ° m m m

kuu • o - ,v-_ _ab ,,-.oco

' I_=_'_) g g o ,--u

-80 -40 o-20 I_ _'O " 040 aoooO_ o ,,-,llO ° I00

Pilot cofl_nlmfiorl (dell)

23
Smith-Geddes Criteria (Research Data)

133 of 207 cases (64%) [82 of 91 PIOs (90%)]

o0oo........4::. o_ OooOoo moo o • PIO

oo°° o o,o _11 ira_

ooq_ o I_"

o o° I'_ ,,,:, •

NoPIO _ oo

°o°° _ >^o J,,4o_ a m m_

oo o_ q_mjodJ= •

%'°I'P"°,%e_, I' "'__8.,. '_'"

(Sotrm ¢1_1 flWI Type II o •

mquwement) Ty_, I Criterion

J Frequency

I,(''<''_)

-_ 40 .1N'_ -120 -140 -180 -180 _ -220 -240 -260

Pit_ Attitude Pha_ AnglemtOtterton Frequ_cy (dig)

0.35

0.3

025

Pile

DIItay, 0.2

(,to)

015

0.1

0.05 -PtO Po_db_

(Pltch RaN

erlhoot)

0

0

Bandwidth Criteria Applied

to Real Airplanes

45 of 49 cases (92%) [20 of 24 PIOs (83%)]

=m_=, D NO PIO I

I

_. • PIO (Not Prlld_) I

•,,i_ _ .... , .,,.. ............

o " #i:r.il_ i-_ "_- "" "-- _i " "

i -FTI .._ ,i =_ _.___ -- '_ III Ji...i

"= NO PIO

2 3 4

Attib_e Ban_twidth, a_ewe (md/mc)

24
Application to Rate-Limited Configurations

Example: Frequency sweeps from LAMARS simulation

(20-deg/sec RL, unstable open-loop; 1 of 5 pilots encountered divergent PtOs)

o25 i plo (Lev= 31 17DVR20 [_ 17OV1_0

C_nSg 17DVR20

AJlitudep_me i-_--

I_llBy ,

015 t

01 PIO _ if

PIto_ Rlm

O'viN'lhQQtJl

EXCUI_e

005

(L_= 2)

oL

o 05

,AL=-=

No PrO

(Ll_e _9

2 2,5

_,t70 . ,_ 1-

(ImJ_l model) . ,/_CO_ 17 (from

/-_ tmclung ran)

_/ PdlotA:

HQR = 3. PtOR = _r

No 1=10(sabble

if Iqch R=e

Exmn_e)

(L,v= 1)

$ $.5 4 45

Application to Rate-Limited Configurations

Example: Config. 2D from HAVE LIMITS TPS Project

(RL on stable bare airplane; no PIOs reported for discrete tracking task)

o.3!

0.25 l

Phase PIO

Delay. 0.2

_e. : PIO Pou, ible

'_' o.15/(,,_.,__ B,._._=,?,o,,,)

0,I i NoPtO

, (if pitchrate |

005l over=hoot II

" / isexeessive) I

0 L_ __J_

0 1

10

I

Max. sweep size: /

o ±lg

• :t2g J

2_'\710

20_ _i60

I I ;,lira

157, 60'_57

Actuato_rfate limit

(deg/se¢)

2

PitchAttitude Ba_dwidlh,¢_BWe(r=d/sec)

25
Application to Rate-Limited Configurations

Example: Config. 2DU from HAVE LIMITS TPS Project

(Unstable open-loop; divergent PIOs for RL of 60 deglsec and below)

0.3

0.25

Phase

D_day. 0.2

0.15

0.1

0.05

0

0

I'

luze:

[] +lg 20 i Max- _ep

I

• _2g I,. 80

PIO 20 _1_1

PIO Pott_l_e No PIO 157

ow_l_oot Actuator m_ limit

Ill excessive) (deg/sec)

2 3

Pi_chAtlJtudeBa_. =BW 0(ra(gsec)

Inappropriate Control/Response

Sensitivity Contributes to PIO

0.5

1t111/1

lannrdl

P_ 0.4¢_1111/11t.i11

¢ott_tml_

(da_=ac=/_)0.3

ozC_,

p_Sl

o::

 IAL..--°

Pitch Example: TIFS Flared Landing Data

4_ 0 7 r

1 a_ 8_id- P_ _

111

p_] PlO

p.4ts_tl

No PIO

0.t 0.2 0.3

26
R4_I o0mrrmn_l

=.en,_Wty,

tdeCV=_,'_,}

80,_ R R R

I 3,2/3/13,3 5/4- .7/6/- -_-

R IndlceteerlNehetrepork_

H_ _ PtW_ _

R (input= =4_reled i_ ¢amm=)

40 _J 2.4/3'2.3 3.4,_44,/6

3O

2o C_ _ e. 0 0 _ • PIO

No PlO

tO

Inappropriate Control/Response

Sensitivity Contributes to PIO

Roll Example: LATHOS (T R = 0.45 sec data)

0.25

Airplane Bandwidth Criteria for Roll

• Much smaller data base

- Not as many real experiences

- Most research experimentsdid not record PIO ratings

• Limits proposed in WL-TR-94-3162:

0.2

Roll

Altitude 0.11

Phase

Delay, o.1,1

(rad/seC)O. 1

N_od mo_ data

In_s crag

Susceptible to PIO

(l..ovel 3)

Scsca_ to P_, (t.evJ 2)

8u=ce_ibOe to PIO If =emll_vity j exc_sive (_ 2)

I

No PIO

0.evel 1)

0

0 1 2

Roll Attitude Bmndwidtfl Fr_luency , meLwo (rad/sec)

27
Recommendations

• Apply criteria as early in development as possible

• Focus especially on Phase Delay limits

- NOgreater than 0.14 sec in pitch or roll

• If feel system dynamics are not known or are known

to be very good, limits excluding feel system are

- No greater than 0.09 sec in pitch or roll

• Use criteria for all amplitudes of control input, up to

maximum possible

- Examine frequency-sweep results if coherence drops

28
PHANTOM WORKS

Stability, Control & Flying Qualities

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

Ken F. Rossitto and Edmund J. Field

Boeing, Long Beach

PIO Workshop

NASA Dryden

April 6-8, 1999

Between 1992 and 1994 The Boeing Company, Long Beach, performed a

series of flying qualities experiments concerning transport aircraft. The

experiments were performed in cooperation with the USAF (focal point Dave

Leggett) and NASA Langley (focal point Bruce Jackson). Both government

partners provided evaluation pilots, the USAF also contributed funding for

flight evaluations.

The purpose of the experiments was to generate a longi tudinal flying qualities

database that could be used for criteria development. The flying qualities

results of these experiments will be presented in a paper at the AIAA

Atmospheric Flight Mechanics conference this August in Portland, Oregon I.

The results of the experiments have also been analyzed to identify PIO

tendencies in the aircraft configurations evaluated. Results from these analyses

will be presented here.

After reviewing the background to the experiments and the approach taken, the

evaluation task will be discussed. The results, as they apply to flying qualities

criteria, will then be presented. Finally, PIO prediction criteria based on the

results will be presented.

1. Field, Edmund J., and Rossitto, Ken R., "Approach and Landing Longitudinal Flying

Qualities for Large Transports Based on In-Flight Results", AIAA-99-4095, presented at the

AIAA Atmospheric Flight Mechanics conference, Portland, Oregon, August 1999.

29
<_ PHANTOM WORKS

$tabllk'y, Conrrot & Flytru Quallt_'s

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

Background

• Requirements for transports not well defined and supported.

• Active control technology make existing flying qualities criteria

obsolete.

Approach

• Develop/validate flying qualities and PIO prediction criteria

and design requirements through a series of generic in-flight

simulation experiments.

Background

Hying qualities requirements for transport aircraft are not well defined and

supported:

• FAILs and JARs are very l im ited

oMilitary speci fi cations are m ore fighter oriented

•Limi ted database on 1 mill ion pound airplanes.

Additional ly, active control technology makes existing flying qualities criteria,

where they exist, obsolete.

Approach

To develop / validate criteria and design requirements through a series of

generi c in-fli ght simulation experiments. Need:

•Preferred response type

•Pitch ax is dy nam ic s

•Pitch axis time delays

3O
PHANTOM WORKS

S|a_l_, Con_"(_ & F_r_ QIwa_tLmll B

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

USAF / Calspan Total In-Flight Simulator (TIFS)

The facility used for the experiment was the USAF Total In-Flight Simulator

(TIFS), operated by Calspan, Buffalo, NY.

Most approaches were fl own into Niagara Airport, though some were fl own at

Buffalo.

31
,_ PHANTOM WORKS

Stab,_ty. Control & Fly_J Qua#ms

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

Offset Approach and Landing Task

_• Simulated touchdowns 1 .... '.....

i " Discrete vertical gusts /

The evaluation task used for the experiment was an offset approach and

landing. The lateral offset of 300 feet was corrected at around 200 feet AGL

and required an additional pitch axis "duck under" to land on the aim point.

Desired performance criteria were:

Touchdown between 1000 and 1500 feet past threshold

Touchdown within 10 feet of centerline

Touchdown sink rate between 0 and 4 feet/second

No PIO

Adequate performance criteria were:

Touchdown between 750 and 2250 feet past threshold

Touchdown within 27 feet of centerline

Touchdown sink rate between 4 and 7 feet/second

All data reported here resulted from simulated landings performed to match

the pilot's correct "eye-height" at the landing point in the simulated aircraft.

32
<_ PHANTOM WORKS

Stability, Conb'ol& Flylrrj Qualities

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

Angle-of-Attack Response-Type Configurations Evaluated

The flying qualities experiment evaluated a range of different dynamics for a

one million pound transport aircraft. The bulk of the data collected was for an

angle-of-attack (or conventional) response-type. Only that data will be

presented here.

Experiment variables were:

n/_x: 2.3 and 3.9

CAP: 0.025, 0.07, 0.2 and 0.6

Time delay: 125,250 and 400 msec

Additionally, two pitch sensitivities were evaluated. The majority of the

evaluations were with a pitch sensitivity of 0.3 deg/s2/lb, and only that data is

presented. A pitch sensitivity of 0.45 deg/s2/lb was also evaluated for selected

configurations.

33
<_ PHANTOM WORKS Criteria for category I PIOs of Transports Based

on Equivalent Systems end Bandwidth

SrabJIl_, Conb_l & FI_ QuaHF_s

Cooper-Harper Ratings (CHRs) Support The CAP Theory

Level 1 / 2 CAP boundary could be raised slightly

looo

I r wl CH_ p,m_ln m* v_ p*ch S_ ol o 3 oe_W.#'_ e

- L

lOaD

oi

1 Io 1oo

nl,_ (B/rid)

BOIJNO

The results for the configurations with zero added time delay (125 msec

baseline configurations) are plotted on the existing Military specification CAP

boundaries. Cooper-Harper ratings for each pilot are presented together with a

"Trendline FQ Level". This trendline flying qualities level was determined

from the individual ratings, the median rating and pilot comments.

Additionally, experimental issues, such as quality of model following in the

TIFS, were assessed. These trendline flying qualities levels have been fixed

and are now used for development of fty ing quali ties cri teria.

The trendline flying qualities levels support the theory behind the CAP

criterion. Additionally they support the raising of the Level 1/2 boundary.

For more details and discussion of these results refer to the AIAA paper

mentioned above.

34
<_ PHANTOM WORKS Criteria for Category l PlOs of Transports Based

on Equivalent Systems and Bandwidth

Sfablllty, Control & Flying Quallf_s I

Cooper-Harper Ratings Show Correlation Between CAP & Time Delay

The results show a multi-parameter correlation between CAP and Time Delay

_ao_

cap

v v

°_100 ol oz 03 o_ os

With the time delay configurations added CAP is plotted against Time Delay.

Note that the two values of n/ct yield slightly different values of CAP, except

for the lowest value of CAP (represented by the circle) which both share the

sam e value.

It is clear from this plot that there is a multi-parameter link between CAP and

Tim e Dela y i n the pi lots' perception o f fly in g quali ties.

35
PHANTOM WORKS Criteria for Category I PIOs of Transports Based

on Equivalent Systems and Bandwidth

StlbJU_/. Conb'oJ _ FJyJrq Qull#t_ls

Correlation of Results with Flying Qualities Criteria

Results do not support MIL-STD requirements

,ooo _ I

CAP 1 ,.

.+]o. 1

v 1 +

D_a

°°1oo o_ o2 ol o,I os

Itqubvlle.tTl_ Dell_l 1_©)

Proposed boundaries fit the data better

looo

1o+ ., ._._

W / ; . +,.'

c+,P

,,,:. i/ +/"

(_',") (T,+_) +_ '; /./ ';

/"

°°10o o

Itm.+*al..I'r+m* O_Imy, (.+c_fL ,_._+LrjAV+,d_,

When the MIL-STD 1797 flying qualities level limit boundaries are added to

the plot of CAP versus time delay (left hand plot) it is clear that these

requirements neither match the data nor allow for the observed multiparameter correlation between CAP and tim e delay.

New flying qualities boundaries have been developed and are proposed (right

hand plot). These boundaries reflect the multi-parameter correlation between

CAP and time delay that were identified from pilot ratings and comments.

These trends have also been observed the results of other ground-based

simulation experiments.

Note: For clarity only the "Trendline Flying Qualities Level" is presented on

all charts from here.

36
<_ PHANTOM WORKS Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

Stabittly, Control & Flyln_ Quall_bs

PiO Boundaries Proposed Based on CAP / LOES Parameters

PlO boundaries reflect the multi-parameter correlation between CAP and Time Delay

_ooo looo

F

i ......................... ._--,°- J L .......................... .-°,.----

• * • ./,'"

...... ?"

o xo-. : _, /,/ ....

"o

°°loo oi oa o:l

CAP

.:,,-....... ;-. ,;,

e le

.

04 0$ 00100 Ol O_ 03 04 OS

E_utv*lehsTim*l ID*ll_ qse¢)

Analysis of the PIO ratings and pilot comments from the experiments led to

the awarding of a "PIO Tendency Classification" to each configuration. This

was achieved in the same way as the earlier "Trendline Flying Qualities

Level". Each configuration was awarded a classification of "No PIO", "PIO

Tendency" or "PIO".

Boundaries delineating the regions of these classifications reflect the same

multi-parameter correlation between CAP and time delay as was observed in

the flying qualities analysis. The limit of "No PIO" boundary appears to be

slightly more relaxed than the Level 1 fimit boundary. This is based upon the

configurations for a CAP of 0.6 and time delay of 250 msec. These

configurations e xhi bi ted only marginal PIO tendency, but sufficient to excl ude

them from classification of "No PIO". Hence the boundary was drawn close to

these configurati ons.

However, the "PIO" fimit boundary appears more stringent than the Level 2

li mi t boundary.

37
_ PHANTOM WORKS

Stability, Control& F_/_ Qua_bs

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

Cooper-Harper Ratings Support The Bandwidth Theory

Level 2 / 3 boundaries could be relaxed significantly

o;s _,,_,_._-.--_.---/--; _ . / ...............

o_ ...._':i::4="_; _ i t"":': ........... :-.--i....

o_ _ ! / 1 " , r !

O0 02 04 06 08 _0 12 14 16 18 20 22 24 26 28 30

PdCh An_ude Bandwi_th,¢l_ (ra_J_oc)

_ ,m'O,l¢.f NO

When the results of the flying qualities experiment are plotted on the

Bandwidth Criterion, it is clear they support the theory of the criterion.

However, they also support the significant relaxation of the Level 2/3

boundary.

38
<_ PHANTOM WORKS

$tablll_, Confrol & Fly;rig Qualities

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

The Data Support the Proposed Bandwidth / PIO Boundaries

The addition of "PIO classification" boundaries might provide more insight

0 50

045

040

035

Z 0 3_

025

020

015

050 L

o45 .... _ - !- a_.z3_o*,d_l_O_o_._o_l_,_ _--_" "'7--

038_ ..................................................

e¢ m '

030_-

-- _ T_':_---_:-_-_T_ "- - -:- --:'- "-_- - t --: -- -

o_8-I _,_'"_"!.--_.---÷--- -_- --_-.- -!- -!-- :---+- -! -i'- ÷-- _....

I

oo0, ; ; , i : ; i ; i i i i

00 02 04 06 08 10 12 14 16 18 20 22 24 28 28 30

P_c_ Allffude Ba_widl_,l_Bw_ _ (radtsec)

When the PIO tendency classifications are plotted on the Bandwidth

requirement they support the boundaries delineating the different PIO

susceptibili ty regions. This may not be i mm ediately obvi ous, but the following

discussion wi 11show this.

The two configurations that were classifi ed "No PIO" fall just above the lower

li mi t of the "Susceptibl e if Flight Path Bandwidth Insufficient" zone. For these

configurations the flight path bandwidth was sufficient, and so they correlate

with the cri teflon.

The configurations with lower bandwidth (the diamonds and triangles) but

nominal 125 msec of time delay all had flight path bandwidths below the

Level 1 limit, and hence are predicted susceptible to PIO. Note that the pitch

sensitivity of the configurations represented by the triangles may have been

high for their pitch dynamics, possibly the cause of the increased PIO

susceptib ili ty of these configurati ons.

All configurations with _:_ greater than 0.15 sec are predicted "Susceptible to

PIO", and these tendenci es were observed during the evaluations.

However, the criterion does not account for degrees of PIO susceptibility, as

does the proposed criterion based on CAP parameters. This could be addressed

by the inclusion of a diagonal line in the "Susceptible to PIO" region,

approximately equidistant from the existing and proposed upper Level 2 limit

on the flying qualities requirement (the plot on the left).

39
PHANTOM WORKS

StabtMy, Control & Flying Q_at/t_s

Criteria for Category I PlOs of Transports Based

on Equivalent Systems and Bandwidth

Conclusions

• Level 1 / 2 CAP boundary could be raised to 0.3

• There is a multi-parameter correlation between CAP and

time delay

• This same correlation is reflected in PIO tendencies

• PIO boundaries were proposed based upon LOES

parameters

• Level 2 / 3 pitch Bandwidth boundary could be relaxed

• The data supports the proposed Bandwidth / PIO criterion

PHANTOM WORKS

Scat_llty. ConU'04 & Flying Qu_,libes

Criteria for Category I PtOs of Transports Based

on Equivalent Systems and Bandwidth

Video of TIFS Landing

Ground View

Pilot View

Configuration:

Angle-of-attack response-type

• n/(x = 3.9 g/rad

• e)'sp = 0.3 rad/sec

Te = 0.125 sec

4O
Designing to Prevent PIO

John C. Gibson

Consultant,

British Aerospace

41
Safety-related PIO

is like the Sword of Damocles, that may:

• break the hair and fall on you if you ignore it,

• but it can also act as a constant reminder

if you act to chain it safely to the ceiling.

• Which one it is depends on you, the

designer

42
0"25

0"20

 ,o.15

0.10

0.05

0

0

Pitch

Level 1

I

1,V;If _Jcntinldyro_attit,_

Roll

Level 1

1(_,_ l(b)

2 3 4 5 6

Attitude bandwidth rad/sec

1(a) has critical damping and low PIO gain,

with translation control qualities that

remain constant as bandwidth reduces

and phase delay increases, while the

attitude control becomes untidy.

1(b) has Level 1 damping (0.5), phase delay

and bandwidth to ADS-33C, but degrades

to dangerous PIO due to high PIO gain and

motion coupling as phase delay increases.

Figure 1 Generic ASTOVL research:

Lateraltranslationhandlingin rollattitudemode

43
PIIch atitude

open loop

gain dB

-180

Pitch atitude .,- --..

open loop phase angle ,"._1 I ""tcleg] -150 -120 ,' I 4¢60

,' I 1_, ;

' _I oIo, '

I O. I

s I ,

, j ' • / I _.'

° i:

m w • _@

1 _".-j// t i ."

L// 7 i -

,---,,,

' ,/^,_ ,_._L.-"" General handling region,

,_' _/" _ "d_v /i°; 4_ Including low order" type PlO

, , ///o.s/

I/"//'_e"-"_ "Good" shape

: "'i/_ lo.e i

,'_ .,"_,'/,t,

' 1 / :Glbson cdterlon" " _ / _1.o ¢ boundaries (Ref. 4)

I

,

I

/"

t

"\

Safety-related "high order" type PlO region

Figure 2 Frequency response qualities illustrated by non-parametric shape

44
[ck_qb]

. 0 , .... -": .I

-2_o -,_o -1so -1_...,r2)- 400 -_ .--.,. g,[c_eol

_r* °" s/

_§ s

;." / s _

,4,' .121 _/

Jll / ............ 1st version pre-filter

,,,I "/ - - - - - - 2nd version pre-filter

_5 r_ 3rd version pre-filter

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

/

I

!

Version

1st

"l'ime to transient

acceleration peak

PIO frequency

Bandwidth

I

m(-z2o) Hz

0.245

2nd 0-245

3rd 0.26

Unaug, 0"145

m(.]8o)Hz t_ sec

0.43 0-35

0"455 0-3

0"535 0"22

0"44 O"12

Pre-filter characteristics

(1 + 0"157S)/(1 + 0"47s)

(1 + 0"133Sy(1 + 0.454s1(1 + 0"0625Sl

(1.0)

(I"0)

Figure 3 Tomado pitch attitude responses at landing: solution to PIO

by development of the command pre-filter.

The unaugmented and third version pre-filtered dynamics are PlO-free.

45
Pilot gain for PIO

is 4 Ib/deg

o-V

/

t

Pilot gain for PIO T

is 25 Ib/deg I

IO(m)l

i |

IF, (m)l dl

-220

-21

1-51

_¢3

-_4o -_oo_pldeg]-eO

I I !

I

0"1 _0"1

i.4 0-5 ..... High order response type:

PIO likely because of

- Low PIO frequency

- High aircraft gain

- Large phase rate

Low order response type:

PIO unlikely because of

f . High PIO frequency

._ - Low aircraft gain

- Low phase rate

ase rate:the local slope of phase lag

vs. frequency at the -180° phase angle point

Severe PIO I

,/- _ Level 3

Phase

rate 200. _/ ,f.___ --"_vel 2

I o\ rate PI

100 .m. 11

O0 0"4 0"8 " 1"2 1"6 -0

m(-z8o) [Hz]

Figure 4 PIO tendency indicators and design guidelines

derived from LAHOS etc.

4(;
le(_,)l

F,--T Id + .)8o._ Phase[deg]

-16o -12o-1-oo\

-4 _ [deg] _ A p

Example of optimum response j ,_ ] m(.,,)

-8_ to resist PIg _".

. A'v_rage phase ra_e

Amplitude boundaries: = {-A phase angle + m(-leo)} deg/Hz

"-" V_ " Upper limits of attitude

._ .20_,,r frequency response in Phase delay z-p

"L1 ..._ the PIg region - {Average phase rate + 720}

y_Y_"-..Level 1" upper hmit

,, for optimum design

(b) Definition of average phase rate

(a) PIg gain limit criterion

Average

phase

rate

[deg/Hz]

25(;

2OO

15C

100

5O

0 _

0

f

L3

L2

L1

deg/Hz

195

145

85

0.2 0.4 0-6 0-8 1.0 1.2 1.4

fz)(.]8o) [Hz]

(C) Phase rate and frequency criterion

(,,-p)

(0.27)

(0.20)

(o.12)

Figure 5 Final development of PIg 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 PIg requires

the more stringent Level 1" gain, phase rate and frequency limits.

47
l e(o))I 81,

_I d. I

[':"°""] 't

0_ 0

" 4 " _ [deg]

L2 //IF ............ 1st version pre-filter

::;_ - . . . . - 2nd version pre-filter

LI" F 3rd version pre-filter

=_ .. m Unaugmented

I

Average

Phase 4

Rate •

[deg/Hz] ]

1 T

50

1st pre-filter f _

I \1/

I !

I /__ 2ndpre-fiiter

t !_-- i i

I deg/Hz

0

0 0.2 0.4 0-6 0.8 1.0 1.2 1.4

o)(-lso)[Hz]

Figure 6 Tomado viewed in retrospect against author's later criteria

Note: although the 3rd pre-filter just satisfies the criterion and

has prevented PIO for 20 years, it would not have been

accepted as a new design by subsequent criteria.

48
Pitch attitude [(leg] ±0.87 • +0.48 •

t= _j,:;

............... ............. ." ..................., L:....._X,2_o_.s_; _-;.-:... -,;_.Tailpiane[ceg];.... _-._.,<,

it-" " ' -"_ .... ' /

!__-. s_ _ I_ __.... _.'_ ..... _

•E;._ s=k_.[,_] ........ ::::, ..... !.

L st_kpumping _ _ Incipient PIO _ Tornado

Pilot attempts to (2nd control law)

freeze the stick

._,.__t_.___._ _ t _..

-T---: ,', i " t_,,?- ..:....:..._.L..L..£_....,....: ...... _.:..... ........... :..-...:.

• I; r ' . PITIItl

• l " _ RAIrI " :

:-4_._ ,._I_...........

• • . - . .

ITIP.,K.

_ Pimp =_I _i

._r : t _- : _-' --i MJ-'-

/':""xl' "_=_ ......... _"."I_'_"" .............

'o.4t. ' : =_,., -n"

"" _IN

Figure 7 Effect of design process on stick pumping and

associated PIO resistance

49
Ii ' ----- Unear

Gain dB :!:5deg

•....-._ ........ _. ._._._'_.;_:._.__ .

_:_+17.5¢leg_ __+5

•.,:,J:...:.3;_ _.:-.'.._;-_F:_,_; _]k._,_...: .:-_ -

•_"_,._:_.."'.; _._"LL.......

0"1 1'0 5"0

Unear

Quasi-linear actuator

response derived from

rig tests

deg

[deg/_b]

"--_ _ [_2]

Demanded /" /- -a__.. ---_',=

/ / - " " ._ -1_ _ _'_llnitial rate limit onset

+12"5_ /" .- _, " "---_ frequencies [Hz]

±_o°" ,, " /:_o_ I

- -0. ,, / IPtO_equencies

+7-5"/-,- .,. -*-*"_" i [HzI[Hz]

,, .," / L:.:_J Unear response

:1:5°°" - / .... Rate saturated

.... rRa_0n se

Figure 8 Significant non-linear actuation effects on PIO characteristics

5O
Attitude dropback ratio = 5.5 sec. . _._6e/

= P'_":_'f

p..._. Pitch rate overshoot ratio 6

ItSll • _,_- ,p

attitude i Path time delay = 0 5 sec. _ ...... " "_'e

angle . _/.,.'_ +,,,..-"""

._I" 0"°- "_ 110 I_3 secs. 2_)

Nominal YF-12 time response

at Mach 3 cruise

I

-5

Pitch . _%_e/

attitude

Flight I _/_ _/_e

angle I / .....,_..,_+x.w"-

+1 +" ,+ secs 2+

Short period roughly approximated by:

O)sp= 0-5 radh_c

_sp =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

51

Session II

53

Replicating HAVE PIO on the

NASA Ames VMS

Jeffery Schroeder

NASA Ames Research Center

Outline

• Introduction

• Experiment description

• Results

• Known simulation/flight disparities

• Conclusions

55
Introduction

• Ground-based simulation has not had much

success in predicting PIOs

• National Research Council recommended

high priority be given to validating

simulation

• Previous fiight-test study (HAVE PIO)

offers a set of pitch data for validation

Introduction

Wright Laboratory replicated in-flight study

using two fixed-base simulators

Purpose of this study:

- Determine if the amount of platform motion

affects ability to replicate in-flight results

56
Experiment description

• Math model

• Task

• Visual system

• Motion configurations

• Safety pilot and miscellany

Experiment description

Math model

Long. r. , I I

stick _ Gearing

position L , , ,oo, ,oo,oH, oot-.

l Pitchrate,ansle-of-attack I

• NT-33 airframe simulated w/stability derivs.

• 18 sets of pitch dynamics

57
Experiment description

Task

i

inn

l/\,\

Three approaches: .......... ---

1. Left offset// 2. Stra Lghtin _ 3. Right offset

_._landing performance

Adequate landing performance

Experiment description

Image system

58
Experiment description

Motion configurations

• Vertical Motion Simulator used to simulate all motion configurations

Vertical Motion Simulator

displacements

i.... Typicalhexapoddisplacements

(5 ft stroke)

t

No motion

ve

/

motion drive logic

Classical motion drive logic

Experiment description

Safety pilot and miscellany

• Automated safety pilot assumed command

if situation deemed hazardous

- Nosegear sink rate > 8 ft/sec when below 12 ft

• Stick ergonomics and force-feel closely

matched aircraft

• Five test pilots (3 NASA, 1 FAA, 1 Boeing)

flew all combinations of motion and aircraft

configurations (randomized)

59
Results

• Example PIO

• Handling qualities ratings

• Pilot confidence ratings

• PIO ratings

• Touchdown velocities

Pitch rate (deg/s)

Example PIO

0.$

Pilot vertical _n__l_ _

_celemtion (g)

i

Time (_)

Large motion satisfactorily simulates pilot normal acceleration

60
Handling qualities ratings

Simulation versus flight

Large motion Small motion No motion

Sim worse li i___ _

than flight

Sim better [.3

than flight

-6

Aircraft Aircraft Aircraft

Configuration Configuration Configuration

Large motion had more ratings within +/- 1 of flight rating

Pilot confidence factors

High

Pilot confidence factor Mot

Min

n=90

t-------

None Small Large

Motion configuration

More confidence in rating with more motion

61
Sim worse _' 2

than flight /

Sim better

than flight

PIO ratings

Simulation versus flight

Large motion Small motion No motion

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

• • .I I l= I

o,.,,,.,., iil-..,p..:.m.!

I ..I

...... |1,.

Aircraft Aircraft Aircraft

configuration configuration configuration

Large motion had more ratings within +/- 1 of flight rating

Touchdown velocities

0

Touchdown

vertical

velocity

(ft/sec)

-2

-4

-6

n=gO

None Small Large

Motion configuration

Large motion allowed better touchdown sink rate control

62
Known simulation/flight disparities

Likely top 5

• Stress-induced environment

• Visual content

• Different evaluation pilots

• Simple automatic versus real safety pilot

• Field-of-view

Conclusions

• With large motion:

- handling qualities ratings correlated best with

flight

- higher pilot confidence ratings achieved

- PIO ratings correlated best with flight

- lower touchdown velocities resulted

• Only large motion provided high fidelity

vertical motion cues

• List of disparities between simulation and

flight suggests future work

63

Replicating HAVE PIO on Air Force Simulators

Ba T. Nguyen, Air Force Research Laboratory

(Report Number 6 is not available for printing at this time)

65

PREDICTION OF LONGITUDINAL PILOTINDUCED OSCILLATIONS USING A LOW

ORDER EQUIVALENT SYSTEM

APPROACH.

John Hodgkinson and Paul T. Glessner

The Boeing Company, Phantom Works, Advanced Transports and Tankers

Long Beach, California

David G. Mitchell

Hoh Aeronautics, Inc.

Lomita, California

Abstract

A study was undertaken to determine whether longitudinal low order

equivalent system parameters could be used to predict pil ot-induced

oscillations (PIOs), also known as adverse aircrafi-pil ot coupling (APC), for

high order aircraft pitch dynamics. The study was confined to linear dynamic

models, and therefore to Category ] PIOs. Variable stability aircraft results

were used from three data sources simulating fighter up-and-away

maneuvering, fighter touchdown, and large transport touchdown. The

equivalent system parameters (alone or in combination) from the current US

Military Standard correlated well with incipient or developed PIOs. Excessive

equivalent time delay was by far the most frequent cause of PIO, and a few

cases were explained by low short period damping, low short period frequency

and low maneuvering stick force gradient. A high-gain asymptote parameter

offered some additional insight into pilot loop closures with large delays.

Hodgkinson, Glessner and Mitchell

67
_. BO_wjAIfO •

Questions

• Can LOES parameters predict PIO?

• IfLOES parameters are good, no PIO?

• IfLOES parameters are bad, can get PIO?

• Do we need dedicated criteria instead?

PIO Prediction using equivalent system criteria

In addition, we would ideally like to answer the questions:

.If the equivalent system parameters were good compared with

the equivalent system criteria, did the pi lots find no PIO

tendency?

.When the pilots experienced a PIO, did one or more equival ent

system parameters predict a PIO?

.A1 so, if it is difficult to obtain a match for a configuration, can

this also suggest PIO susceptibility?

We were able to answer all these questions to varying degrees.

Hodgkinson, Glessner and Mitchell

68
..._mOf/ND"

PIO Rating (PIOR) Scale

m

_,mmai_ _ plaq _

p_ ,,_'_ .,_ _'_,_

s

PIO ratings awarded by the pilots aided this study.

Hodgkinson, Glessner and Mitchell

69
Three data sources

• Neal-Smith

• LAHOS

• GLT

Correlati on database

Three data sources were utilized. All were from in-f/ight simulations.

Reference 6, Neal and Smith's study, examined up-and-away dynamics of

fighter aircraft. Reference 10, the so-called LAHOS study, considered fighter

dynami cs in the landing approach. The Generi c Large Transport (GLT) study

of Reference 11 was for landing and touchdown dynamics of very large

(approximately 1-million-pound) transports. In these data bases, the pilot

ratings and comments were used to separate the configurations into those

without PIO tendencies, those with incipient PIOs, and those with actual PIOs.

(for Reference definition, see the last two charts, or AIAA Paper 99-

4008,'Prediction of Longi tudinal Pil ot-Induced Osci ]1ations using a Low Order

Equivalent System Approach', John Hodgkinson and Paul T. Glessner, The

Boeing Company, Phantom Works, Advanced Transports and Tankers, Long

Beach, California, and David G. Mitchell, Hoh Aeronautics, Inc., Lomita,

California).

Hodgkinson, Glessner and Mitchell

7O
Ja,_JNO °

LOES form for pitch rate control

K o

(s + L_ )e -_s

2

[s 2 + 2_pCO,,s,S + _,,_,, ]

The accepted method for determining the longitudinal short period equivalent

system is to match the pitch and normal load factor dynamics (at the

in stantaneous center of rotation) simultaneously. Similar parameters are

obtained by matching the pitch rate dynamics alone with the transfer function

shown in the chart, with fixed at the value for the aircraft. The transfer

function numerator includes a gain; the dimensional lift curve slope of the

aircraft; and a time delay. The denominator includes the short period damping

and undamped natural frequency. For these pitch dynamics, good and bad

values of the parameters are all defined directly or in combination by the

current specification, Reference 1.

Hodgkinson, Glessner and Mitchell

71
Candidate equivalent parameters

• Time delay

• Short period frequency

• Dimensional lift curve slope

• Short period damping

• Stick force per g

• High Gain Asymptote Parameter (HGAP)

Early equivalent systems researchers quickly found that the high frequency

phase lag, or rolloff, of some high order responses was greater than that which

the low order forms could accommodate. Therefore a time delay term was

added to the low order forms. The delay itself eventually became a criterion

for handling quali ties specification (see Reference 1). The High Gain

Asymptote Parameter suggests that a tight pitch 1oop closure by the pilot could

cause unstable pitch oscillations. ( Ashkenas et al Reference 9). Low values of

short period frequency produce sluggi sh dynamics and a low Control

Anticipation Parameter (CAP). Low values of short period damping produce

open-loop oscillations. Combined low stick force per g and low damping

produces dynamic sensi tivi ty. High steady-state sensitivity of response to stick

command can produce PIO, as can combinations of rapid short period

frequency with significant pitch delay. Too-abrupt (too-high) short period

frequency can cause PIO. Fundam entail y conventional aircraft with high

mismatch, i.e., whose dynamics cannot be matched with a conventional

transfer function, are unlikely to have good handling qualities. However, first,

configurations with high mismatches tend to have extreme and unsatisfactory

equivalent parameters, and second, if an inappropriate equivalent system form

is used for an unconventional response-type (like an attitude command

system), then the resulting high mismatch is just a consequence of misuse of

the method.

Hodgkinson, GI essner and Mitchell

72
,...BO_L'-j/led -

Low CAP=PIO for transports

PIOR

6

5

4

3

2

1

i

i

0 0.2 0.4 0.6 0.8

Control Anticipation Parameter, CAP

Control Anticipation parameter (CAP)

Sluggish short period frequency would be expected to correlate with PIO

tendency. When all the CAP data from the experiments were plotted without

regard to other parameters, a tendency to support this expectation emerged, as

seen in this Table:

CAP

Data Source

Neal-Smith

LAHOS

GLT

Apparent tendency for PIO if CAP is less than:

0.2

0.18

0.18

However, further examination of the data shows considerable influence of

other parameters. For example, the low-CAP configurations in the Neal-Smith

data generally had high equivalent delays. Thi s is a natural consequence of

how Neal and Smith added lags to fundamentally conventional dynamics to

create their sluggish configurations. Lags not only add equival ent tim e delay at

higher frequencies, but also depress the short period equivalent frequency in

the mi d-frequency range. When the effects of other parameters are separated

from the data, we were let_ with only the GLT data giving a significant

indication of PIO tendency due to low CAP values, as seen in the chart.

Hodgkinson, Glessner and Mitchell

73
,8,O,¢¢-,/AFD °

HGAP:PIO if | / To: > 2g ,p(0,,,,

High Gain

Asymptote

Parameter

_1/7;) = 2g ,re)..,

1/_) =0/ 1/7_, =x

"* [l l'[/ '[ "

1/ 7;), ._["

_.I::L_ [ ">R

High Gain Asymptote Parameter (HGAP)

The early equivalent systems analysis of the Neal-Sm ith data did show a high

correlation of the high gain asymptote parameter with poor ratings (Reference

2) but equivalent time delay, i.e., high frequency phase lag, dominated the

PIO-prone cases. Low values of HGAP would be expected to correlate with

PIO tendency. In the original theory, it was pointed out that an adverse

constellation of roots for the pitch rate transfer function was unlikely for

conventional aircraft, and that additional phase lags (i.e., equivalent delays)

would be needed to cause PIO. Use of the 'free L-alpha' data promised to be a

way of incorporating some lag into the basic root array by shifting the lead due

to to artificially high frequencies. That technique also created negative

values of HGAP, correlating with PIO. However, since freeing in the

matching process is quite artificial, and the resulting delay values are not

comparable with most studies, we do not present these data here.

Hodgkinson, Glessner and Mitchell

74
Low HGAP=PIO for Neal-Smith

I

u

.Q

Ntis.StrUm

L_ F4.*a J. colts S_**_

* .• e _o

{3OO - . ,D

H_g_ GJm Asymp¢at_ P_ram_ - HGAP

Plotting the HGAP (with fixed L-alpha) against PIO rating for the Neal-Smith

data does show a general trend of worsening rating with smaller HGAP but for

the other data bases the data did not show a clear correlation.

75

Hodgkinson, Glessner and Mitchell
)._,8'aflAgD"

HGAP and equivalent delay...

can HGAP help bad delays?

i[ :::: i

_ r

,. iI

T_ Del_ ITau) - _¢ondt

Plotting HGAP versus time delay for fixed shows that Neal and Smith's

configurations with high time delay in generhl also had low (theoretical ly bad)

values of HGAP. There is a weak suggestion in the right eight data points in

this Figure that the PIO tendency of configurations wi th high delays might be

ame liorated by increasing HGAP.

Hodgkinson, Glessner and Mitchell

76
...._,_B'OfJ, dlCO"

Can HGAP help bad delays in LAHOS

too?

ko Field M CO*tl Iho_m

i

_. , ° o°

Timo Oilily {Tau) - eq_condl

1

I

, Jc,

The LAHOS data also contain this weak suggestion in the region where time

delay is between 0.15 and 0.2. The data are not conclusive enough to suggest

an actual requirement involving H GAP. Further systemati c data involving

HGAP variations are needed.

Hodgkinson, Glessner and Mitchell

77
..._BB_'/A¢O °

Delays cause PIOs (Neal-Smith)

g

i_leaI-Smit h data

_k_ M La Fl:ed U C_s ShOWn

Time Delay (Tau) - leconds

Equivalent time delay

Correlation of this parameter with PIO susceptibility has previously been

noted by researchers including Neal and Smith (Reference 6) and Hodgkinson

et al (Reference 2). Our re-examination of the Neal-Smith data did confirm

the progressive increase in PIO susceptibility with increased delay. The other

data bases allowed only an i ndi cation of when tendencies towards PIO could

be expected. The following Table summarizes the del ay values:

Equivalent Delay

Data Source Tendency for PIO if delay exceeds:

Neal-Smi th 0.12

LAHOS 0.16

GLT 0.25

Definite PIO if

delay exceeds:

0.18

78

Hodgkinson, Glessner and Mitchell
NOE'PNO"

Conclusions

• LOES parameters predict PIOs reliably

• Data bases mostly delay-dominated

• Low CAP for transports causes PIO

• Low Fs/n caused one PIO in Neal-Smith

• HGAP- intriguing interaction with delay?

Conclusions

Short-period equivalent system parameters offer many clue s to longi tudina 1

PIO susceptibility. In the data examined, excessive equivalent time delay was

the chief culprit. For example, in the Neal-Smi th data, every configuration

with a delay exceeding 0.116 seconds had a tendency to PIO. Other

parameters correlating with PIO tendency included low equivalent damping

ratio and low stick force per 'g' for the fighter configurations, and low

equivalent frequency for the transport.

These results suggest that meeting the military equivalent system requirements

would help to avoid PIOs.

The linear parameters used in most of the alternative PIO criteria and in the

equivalent system parameters in this paper evidently address only a part of the

PIO problem. Future work needs to address the roles ofnon-linearities and of

structural dynamics.

Finally, the High Gain Asymptote Parameter (HGAP), based on li near

equivalent system parameters, shows some correlation with PIOs, and there is

some evidence that configurations with marginal equivalent delays may

benefit from larger values of HGAP.

The work in this paper was supported by Hoh Aeronautics, Inc. under their Air

Force Research Laboratory contract on PIOs, and by the Boeing Company.

Hodgkinson, Glessner and Mitchell

79
BgfJ/I¢O °

References

1. Anon, MIL STD 1797, Flying Qualities of Piloted Aircraft, MIL-Prime

Standard and Handbook.

2. Hodgkinson, J, LaManna, W.J., and Heyde, J.L., "Handling Qualities of

Aircraft with Stability and Control Augmentation Systems _ A Fundamental

Approach." J.R.Ae.S., February 1976.

3. Hob, R. H., Mitchell, D.G., and Hodgkinson, J.; "Bandwidth- a Criterion for

Highly Augmented Airplanes". AGARD Conference Proceedings No. 333,

Symposium on Criteria for Handling Qualities of Military Aircraft, Fort

Worth, Texas, US, 19-22 April 1982.

4. Smith, R., and Geddes, N., "Handling Quality Requirements for Advanced

Aircraft Design : Longitudinal Mode". AFFDL-TR-78-154

5. G_son, J. C. "Development of a Methodology For Excellence in Handling

Qualities Design for Fly By Wire Aircraft". Delft University Press, 1999.

References, concluded

6. Neal, P.T., and Smith, R.E., "'An In-Flight Investigation to Develop Control

System Design Criteria for Fighter Airplanes". AFFDL-TR-70-74, December

1970

7. Hodgkinson, J.; Aircraft Handling Qualities, AIAA Education Series, 1999.

8. McRuer, D.T., Ashkenas, I. L., and Graham, D.; Aircraft Dynamics and

Automatic Control, Princeton University Press, Princeton, New Jersey, 1973.

9. Ashkenas, l.L., Jex, H.R., McRuer, D.T.,"Pilot-induced Oscillations: Their

Cause and Analysis". NORAIR report NOR-64-143, July 1954

10. Smith, R.E., "Effects Of Control System Dynamics on Fighter Approach and

Landing Longitudinal Flying Qualities." AFFDL-TR-78-122, March 1978

1 I. Field, E.J., and Rossitto, K.F.; "Approach and Landing Longitudinal Flying

Qualities for Transports Based on In-Flight Results" AIAA Paper 99-4095,

AIAA Atmospheric Flight Mechanics Conference, 9-11 August 1999,

Portland, Oregon, USA

Hodgkinson, Glessner and Mitchell

80
Recommendations to Improve

Future PIO Simulations

Brian Stadler

AFRL/VACD

2180 Eighth St, Suite I

• _ Bldg. 145 Area B

Wright-Part AFB, OH 45433

Phone: (937)255-6526

Fax: (937) 255-9746

E-Mail: Bdan,Stadler@va.afr.af.mil

Why Important?

• Manned simulation is being relied upon ever more

• Virtual Combat Simulations

- Used to design and set aircraft system requirements

- Determine force mixes

• Simulation during aircraft development

- Assess vehicle and train pilots before flight

- Considered alternative to flight test!

• Classic use of simulation (control design tool)

- Assess aircraft handling qualities

- Iterate flight control design with pilot-in-loop

• Modeling and Simulation is perceived as a means to

reduce costs!!

81
PIO S,mulation D iemma

• Historically PIOs not readily uncovered during simulation

experiments

• Often found in flight test and then repeated in simulator

• Several types of PIO initiated for different reasons

- Category h PIOs by linear phenomena, phase loss,

• Empirical Cnteria Exist

• Correlates to bad handling qualities

- Category Ih PIOs caused by non-linear phenomena, rate limiting

position limiting, gradient breaks

• Criteria under development

- Category II1: PIOs caused by mode switching

• PIOs generally occur when pilot is hi qh .(:lain and working

hard at a precision task.

PIO Simulation Background

• AFRLNA PIO Simulation Objectives:

- Attempt to determine reasons why ground based simulations do

not readily uncover PIOs dudng development

- Use a known flight-test truth model to conduct comparisons to

ground based implementation

- Attempt to develop a methodology to uncover potential PIOs in

aircraft more reliably via simulation

• Two truth models:

- HAVE PIO: USAFTPS-TR-85B-S4

- HAVE LIMITS: AFFTC-TR-97-12

• Want simulations to correlate better with flight test

- What do we mean by correlate?

82
Simulation Facilities Used

Mission Simulator l (MS-l)

• Fixed Base, 40Ft Dome

• Me Fadde n Feel System

•Wrap around visuals

• HUD projected

Large Amplitude Multi- Mode

Aerospace Research Simulator

(LAMA RS)

• 5-DOF Simulator

• McFadden Feel Sy stem

•20ft Diameter Sphere on end

of 30 ft beam

•Wrap around visuals

HAVE PIO Phase 1Tests

• HAVE PIO Phase 1 Tests

- Eighteen different configurations

- Linear sources of PIO

- LAMARS (w/wo motion) and MS-1

- Power approach task only

- Priority on replicating NT-33 tests as accurately as possible

83
•::i{:iii, i:_i:_!_(!i!  ̧ :!!i!ii:!_ilqlt_i,lii_!'ii:!i:i_ i/i,_ii, _::ii ..... _ ::_

HAVE PIO Phase 2 Test

Pylons were added to the landing

task to force pilots to fly a

particular path and to hi-light the

touchdown point. Left, Right, and

Centerline Pylons sets were used.

2,5 dog Glide Slope

HAVE PIO Phase 2 Tests

.MS-I

•Power approach only

•Assessed simulation tweaks

•Stick Gain

•Time delay

-Winds/Turb/Gusts

.Pylons

HAVE LIMITS Tests

HAVE LIMITS Tests

- LAMARS with motion (retune)

- SOS and Calspan Discrete task

- Attempt to correlate with NT-33 Test

- Core of an expanded database

- Changed HUD Symbology from NT-33

Error

_. _-.'_ ._ _s_

Aur_d "V 30_ _ - '_ Ahlt_

- 14,5

: --6--

FhghT

Path _ _ Ileading

M,lrkcr 05

I ,_._: ,. .......

t'"

.-

°.

84
Results

• HAVE PIO

- Able to generate Category I PJOs in simulation

- Desired correlation between flight and simulator per configuration

not achieved

- Data trend: good was good, but bad was not as bad

• HAVE LIMITS

- Initial tests uncovered problems with model replication between

what occurred in-flight and what was integrated on simulator

- Category II PIOs replicated in simulation

• Wanted direct correlation with flight test for each

configuration or predictable variation across CooperHarper and PIO Rating Scales

Reason for Differences

• Fundamental difference between handling qualities

evaluations and PIO experiment

- Evaluating a configuration versus searching for defects

• Pilot variability even a larger factor in PIO experiments

- Large variations not unusual

- 3 Pilots do not a make a sufficient sample space

- Pilot technique

• Briefing Techniques

- This has an effect: Reviewing PIO charts, definitions

• Task Definitions

- Already difficult to match reality

• It's a simulation!!!!!!!!

85
" PIO Testing

• Hypothesis: Fundamentally different from standard

handling qualities testing

• During HQ testing pilots are rating the configuration as is,

not actively looking for deficiency

- If we run into PIO great, if not, no PIO

- This does not imply conf_juration is not PIO proof

• PIO requires an active search

• Test matrix and task development require much more

attention and care

• Need real-time measure of pilot effectiveness during task

to keep honest (RMS, Touchdown dispersions)

' Task Generation •

• PIO Testing requires closed loop high gain tasks that

stress pilot/vehicle system

• Approach Task Too Open Loop

- Suggest use of pylons, ILS needles

- Measure pilot performance along path

- If pilot doesn't land is that a CH 10???!!!

• Discrete Tracking Task

- Works well in simulator

- Pilots game system so variations must be used to avoid learning

- Requires Tuning, we found pilots could trip into PIOs especially in

one region!

• Remember: It's a simulation

86
, Tracking Task

,_ Gotcha Region

I .

Pilot had rated this pitch

configuration (2DUR30) in

earlier runs as a CH-2

PIOR-1, During this run a

rate limited roll was added

to increase workload.

" :Pilots

• Natural variability puts pressure on other parts of PIO test

- Need more than 3 pilots, but not just for statistics

- High/Low Gain, Golden Arm, The guy who hates simulators

• Shouldn't fly more than an hour !

- Fatigued pilots good for PIO generation but bad evaluators

- Fresh pilots make good evaluators but poor PIO generators

- When pilots refer more and more to previous runs, break!!!

• Need to keep aggressive by any means necessary

- RMS feedback worked well, but when do we give to pilot?

• Need to reset pilots often

- Good->Bad, follow really bad config with a good config

8?
....Pilot Brieflno .....

• Critical to success of any test.

- Not all Test Pilots have seen a PIO

• Define PIO

- What is a bobble? What is an oscilla_on? Overshoot?

- Does backing out of loop imply PIO and what to do?

• Define tolerable/intolerable workloads and define

adequate and desired.

- Some pilots definitely have a distinct definition of these.

• Pilot ratings in a simulator

- Level 1 ratings reserved, psychological block

- Some pilots won't even give a CH-IO!_!

- Pilot can crash in a plane but not in a simulator

Simulation Motion

• Motion versus no-motion

- Well tuned motion helps

- Extra cueing to pilot, especially of AZ phasing

- Give hint to pilot if something is not right

• Lack of motion puts pilot reliance on visual cueing

- Hard to discern rates of descent

- Visual detail limitations

- During air-to-air tracking scenery isn't important anyway

• Hard to determine value due to interpilot/intrapUot variability

- Can't really determine worth via Cooper Harper Ratings

- Pilot comments have been extremely positive

• If good motion doesn't help does bad motion really

hinder?

88
Motion Work

Objective: Maximize Acceleration Recovery

Use the most motion travel w/o hitting limits

Minimize False cues with proper phasing

" - .... AZ Aircraft

¢

_.a i AZ Recovered

Tirr_ Time Timc

SMTD Washout Tuned Linear Washouts New Non-Linear Drive

Non-Linear: Uses Fuzzy Logic Approach

Uses Predetermined Braking and Retum Profiles

Uses Human Thresholds and Indifference Levels

Wrap Up

Simulation _=Replication!![!!

- Attempting to replicate flight test results dubious effort

• PIO simulations require extra effort in other areas

- Not asking do you like this or not?

- Asking, did you find a problem

• The more pilots the better

• Test setup and pilot brief can do more to trash results

than simulation artifacts

• Task design critical. Can only do so much to simulator

• Motion use recommended, but must be properly tuned to

be of benefit

89
Analyti cal Time Delay

 Measurements

Pilot

O- 16.67mse¢

DiD

_(:)

2reset

Total: C4 77-110msec

SG 52-119 msec

D/D 0-16.67reset 75reset

0-33 3msec 50-66.7msec

HUD

Visual

Display

VLsual

U,(Z) Display

,,VD D/D

vm(s) .__(z)

A/D D/D

0-reset Motion

System

50msec

Measured Time Delays

• Two types of delay measurements in simulators

Time Domain: time to wiggle to time to response

Frequency Domain: Sum-of-Sines phase delay

LAMARS freq domain tests accomplished on motion while both freq and

time measurements were done on visual

_- MS-1 only time domain tests were done on visual

• LAMARS Measured Visual System Delays

- Compuscene transport delay: TD=88msec

- Compuscene End-to-End: TD=108-124msec FD=72msec

- HUD End-to-End: TD=69-153msec

• MS-1 Measured Visual System Delays Time Domain

- Compuscene transport delay: TD=75msec

- Compuscene End-to-End: TD=94-111msec

- HUD End-to-End: TD=69-153msec

9O
Tracking Task

Motion Work

6 Video Channel_. _..__ 5-DOF Cab pitch, roll, yaw,

t-_ _ heave, and sway

_ . ._j .

Pilot Station Sensor Package:

2 Accelerometers Az, Ay

3 Rate Gyros

• Conducted parameter identification of all servo-axes.

• Developed new beam compensation terms.

• Retuned linear washout terms.

- Used new terms during HAVE LIMIT testing

• Non-linear washout scheme developed for AZ cueing

- Implemented tested using Capt. Chapa as test subject

- Initial feedback good both subjective and analytical

91

Session III

93

FAA'S HISTORY

WITH APC

Guy C. Thiel, FAA

FAA'S HISTORY WITH APC

• BACKGROUND

• INITITAL DEVELOPMENT OF CRITERIA

. FINAL CRITERIA & RATINGS SCALE

95
BACKGROUND

• 1993 - Special Certification Review

- High Altitude Turbulence Upsets

• 1994 - Initial Draft Criteria - FBW Program

• 1995 - First Meeting of NRC Committee

• 1996 - New AC 25-7 with APC included

• 1997 - Final Release of AC with Comments

BACKGROUND

• MD-11 INCIDENTS

* FLYING QUALITY RULES

ONLY CLOSED LOOP

NO HIGH ALTITUDE TASKS

96
INCIDENTS

. MD-11 HIGH ALTITUDE UPSETS

• OTHER INCIDENTS

• CAUSES

Basic Handling Qualities ??

Lack of Training

Unusual Atmospheric Conditions

FLYING QUALITY RULES

• Normally Open Loop Tests

• Tasks are not Used in Certification

• High Altitude Flying - Autopilot

97
CRITERIA

• REGULATORY BASIS - FAR 25.1143

• A) The Aircraft must be safely controllable and

maneuverable throughout the flight envelope.

B) Must be possible to make smooth transitions from

one flight condition to other flight conditions without

1) exeeptonai pilot skill, alertness, or strength

2) exceeding airplane limiting load factor

CRITERIA

• Link FAR 25.143

• Handling Qualities Rating Scales FBW AircraR

• FAA Rating Criteria

• Develop APC/PIO Rating Scale

98
IMPLEMENT CRITERIA

• Use Advisory Circular Method

A) New Rules - 5 to 7 Yrs.

B) Add to Flight Test Guide (25-7)

C) Para. for FAR 25.143

• Add Required Maneuvers

• Tie APC Ratings to HQR Section

IMPLEMENT CRITERIA

• Issued Drait of AC 25 - 7 in Early 1996

• Basis for Certification

• Aircratk Tested - MD-11, B-777, IL-96T, A330-

200, Citation X, G-5, Global Express

99
NEW CRITERIA

• Published AC 25 - 7 (Original Criteria)

• Train FAA Test Pilots

Modify Original AC 25-7 Material

TRAIN TEST PILOTS

• Select First Group for Calspan Training

• Interim use of Intitial Group

• Plan for Remaining Pilots

lOO
MODIFY APC CRITERIA

• Because of Results from Past Programs

• Add Operational Maneuvers

• Require Tracking Device

• Modify APC/PIO Rating Scale

101
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102
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130

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103
CALSPAN

_ Ctw_m _ VeWm Tools for the Evaluation Pilots

II

Sum of Sines Tracking Task

(similar in roll)

CA/SPAN

Tools for the Evaluation Pilots

I

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...... _...... !...... ;...... i...... ; ...... i...... ;,...... ;......

: _ : ! ! i

• • m m I m _ m

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Discrete Tracking Task

104
Figure 20-1 FAA APC Rating System

F/U_ Draft 5

12/18/98

_l-- . _ .._11%t _

imr " --_

105

APC/PIO Workshop

NASA Dryden Flight Research Centre

Edwards, California

6-8 April 1999

Graham Weightman, JAA (UK CAA)

APC/PIO Workshop

Dryden Flight Research Centre, 6-8 April 1999

• Initial discussions with FAA in the JAA Flight Study

Group (FSG) on proposed APC text for draft revision to

FAA Flight Test Guide (AC 25-7X) beginning early in

1996

• JAA submitted comments on AC 25-7X (September 1996)

• Further discussions on APC in FSG (reference Flight

Working Paper 599 prepared by FAA)

• JAA has reserved the APC text for the first issue of the

JAA Flight Test Guide (based on AC 25-7A and to be

published for comment shortly) pending further work

107
APC/PIO Workshop

Dryden Flight Research Centre, 6-8 April 1999

• FSG established an ad-hoc Sub-Group to work with FAA

on harmonised guidance material for APC

• FAA (Mel Rogers) invited to chair Sub-Group

• First "kick-off' meeting in Braunschweig, Germany in

January 1999. CAA, LBA, DGAC/CEV, FAA,

A6rospatiale, Airbus and Boeing/AIA present

• Intention to work largely by E-mail

• Target: Draft revision of FWP 599 by June 1999

108
-PIO Flight Te ......

109
Introduction and Disclaimer

This presentation represents a snapshot in

time with regard to Boeing's flight test

experience with Pilot-Induced Oscillations.

The information contained herein is

presented in the hope that in sharing

technical information, safety can be

enhanced through cooperative focus of

research, and reduced duplication of efforts.

110
Agenda

Boeing Flight Test Ew

- Aircraft Scope

- Data Collected

- Maneuvers Used

Need for further work

- Controller Characteristics

- Nonlinearities in Response

- Pilot Aggressiveness

This presentation consists of two parts.

The first is intended to let the technical community know about Boeing

(Commercial) flight test activity with respect to PIO. The scope of aircraft

model s tested, the kinds of data collected, and experi ence regarding vailous

speci fi c eval uation maneuvers will be disc ussed.

The second part of the presentation contains suggestions for focus areas in which

the current state of analytical techniques is not adequate to address many very

real situations which arise in the testing of large commercial jet transport aircraft.

111
PIO Testing History at Boeing

Specific Evaluations carried out since 1995

- 777-200 737-700

- 777-300 737-800

- 757-200 757-300

• Plan to include other models at "windows of

opportunity"

Boeing Commercial Airplanes takes Pil ot Induced Oscillations very seriously

and endeavors to understand the phenomenon to insure that its products do not

exhibit these adverse characteristics. Since 1995, Boeing has undertaken to

evaluate a number of airplane models, and have a plan in place to evaluate others

as opportunities present themselves.

As can be imagined, fully instrumented airplanes are not always easy to come by,

so data is acquired whenever it is available.

112
Intent of Generic Test Program

• Evaluate Each Boeing Airplane Model

• Collect Data

- End-to-End Open Loop Dynamic Response

- Control System Response

- Qualitative Evaluation During High Gain Tasks

- Quantitative Evaluation During High Gain

Tasks

• Document Lessons in Design Requirements

At the outset, Boeing conceived a generic test program which had the intent to

conduct specific evaluations for PIO tendenci es on each Boeing airplane model.

These evaluations were multi-faceted and intended to acquire four different types

of data. These included:

• end-to-end open loop dynamic response

•conrol system response data

oquali tative evaluation during high gain tasks

•quantitative evaluation during high gain tasks

In addition to collecting the data, the results of the testing and subsequent

analysis would be documented as lessons learned in internal design requirements.

113
Maneuvers Flown

Maneuver Flight Condition I Configuration

• FrequencySweeps -_ )i HighAltitudeCruise

• Control Doublets \ Low Altitude Cruise

• Control Releases \ Approach

• Close Formation

• Constant Altitude flybys I

• Lateral S Turns

• Vertical S-Maneuvers (, • Landing

• Offset Landings

The primary maneuvers in the generic plan are shown on the chart

Open loop airplane and control system response data and the qualitative close

tracking task (formation flying) is collected at high and low altitude cruise,

approach, and landing conditions. The runway work is done only in the landing

configuration.

Open loop response data collection, consisting of frequency sweeps, control

doublets, and control releases are self explanatory, and not described further.

114
Formation Flying- Box

20 Feet

_10 Feet

20 Feet 20Feet_

Hold at Corners

for 20 Seconds

10 Feet

10 Feet

A number of speci fi c maneuvers have been used as close tracking tasks in up and

away flight. One of the most effective has been close formation flying. A

particular difficulty in implementation of this technique is that it is mostly

qualitative in nature. Accurate measures of pilot-in-the-loop performance and

and ways to adequately feed it back to the pilot have not been identified.

Although discus sions of over-the-shoulder cameras, heads-up di splays, and

differential GPS installations have taken place, none have as yet been

im pie men ted.

One maneuver used as a piloting task is the formation box maneuver, shown

here. Once the pilot is establi shed in a close refueling position (thought of as the

center of the box), the pilot is asked to rapidly and aggressively acquire a new

position 10 feet to the right. This new position is to be held as closely as possible

for 20 seconds at which time the pilot is asked to acquire a new position 20 feet

below the last. This is similarly held for 20 seconds. The maneuver proceeds

around the "box". This maneuver combines a gross acquisition task with close

tracking in a very high gain environment, and combines both longitudinal and

lateral-directional axes.

The inset shows flying this maneuver with a 777-300 flying against another 777-

300.

115
Formation Flying - Cross

Lead Aircraft

10 Feet

Hold at Ends

for 20 Seconds

A second maneuver used is the formation cross maneuver. Execution of this

maneuver is similar to that for the box.

One element which makes these maneuvers interesting in flight is that the trail

airplane is flying in a curved flowfield. What this means is that to hold at the

lateral ends of the cross requires flying in sideslip, which adds to pilot workload.

The inset shows this maneuver being flown in a 777-200 against a 747-400.

116
Trail Position- Flaps Down

When transitioning to the approach and landing configurations, the lead aircraft

also transitions in order to match fli ght speeds. Shown here, the trail pilot is

looking rather directly at the upper surfaces of the very large triple slotted flaps

of the leading 747.

Now while the vertical tail of the trail airplane is certainly i mm ersed in the wake

of the lead airplane in all conditions--and the buffet is noticable--the wake grows

considerably for these flap down conditions. This increased the workload for the

777 airplanes, but the attendant buffeting was simply unacceptable for the

shorter, lighter 737 airplanes. The task was not possible given the severity of the

buffeting for that (737) airplane. So the entire task was moved to the wingtip of

the lead airplane.

117
Formation Flying - Wingtip

Maneuvers

_)J Also: Altitude Up and Down

I 20 Feet 20 Feet Maintaining

Lateral and Longitudinal Position

While the wingtip formation maneuvers were planned for all airplanes anyway, it

was discovered that this was the only practical position to evaluate the flaps

down conditions for the 737.

The wingtip maneuvers are shown here, including transitions fore and aft, in and

out, and up and down. In addition the trail airplane was asked to follow the lead

through turning maneuvers, keeping station on the wing tip.

These maneuvers proved to be very demanding. Compared to the refueling

position, the wingtip position provided a much s mal ler target (the wing tip itself),

which the pilot could see with better precision, and the target was much more

active. Especially as the leader turned, the wingtip moved around significantly,

generating a very demanding tracking task.

The inset shows a 777-200 flying against the 747-400 in the wingtip position.

The evaluation pi lot is focused very intently on what the lead aircraft is doing.

The situation is just as dramatic when vi ewed from the lead aircraft.

118
Close Wingtip Position

This is a 737-700 being flown against a 737-800. The distances are short, and

pilot gain is very high.

119
Formation Flying Summary

• Single Highest Gain Task

• Maneuvers Combine Acquisition with Tracking

• Learned Task Requiring Experience

• Wingtip Tracking Probably Most Effective

• Difficult to Measure Performance (and Feed Back to

Pilots)

- DGPS in the Future?

• Difficult to Enforce Performance Requirements

• Difficult to Get Consistent Level of Aggressiveness

To summarize Boeing experience with close formation fly ing as a man euver to

explore APC tendencies, it can be said that it provides a very high gain task

which combines gross acquisition with tight tracking.

At the same time, it is very di fficult to measure the pil ot/vehicl e performance and

feed that back to the pilot in a meaningful, quantitative way. In addition, and

perhaps because of the lack of performance information, it is very difficult to

achieve consistency in aggressiveness across several evaluation pilots.

120
Constant Altitude Flyby

Intended to "Extend" the Flare for Analysis

Involves both Acquisition and Tracking

- Fly ILS to 50 Feet

- Flare and Maintain 50 +/- 10 Feet for Length of

Runway

- Maintain Centerline

- PNF Calls Radar Altitude

Another set of maneuvers used to explore APC tendencies has involved flying

close to the runway. Original ly, the fl yb y task was conceived to provide in sigh t

into the pilot/vehicle combination in the flare. Upon examination, if done

properly, a flare maneuver takes only a few seconds. On large transports with

natural frequencies on the same order, it is difficult to gain much understanding

about the interaction. So this maneuver was conceived to provide an extended

time period for data gathering. The maneuver involves acquisition and tracking

in a high precision environment.

The pilot is asked to flare and maintain 50 +/- 10 feet for the length of the

runway. Typically, the pilot will close a loop around radar altitude, with the pi lot

not flying calling radar altitude continuously. During the maneuver, the pilot is

asked to maintain the runway centerli ne.

It was discovered that the most difficult part of the task was making the power

adjustment in the round-out. Too little power and airspeed would bleed away in

the level segment; too much, and the airplane would accelerate or climb.

Pilots descried the task as challenging but not impossible.

121
Flight Performance

Pilots Characterized Task as "Demanding, but not

Impossible"

Power Setting in Flare Requires Precision

i

..°

.,°

\ I

/

., °; .. ...... ::... °........

An example time history shows that the desired performance level could be met.

It is interesting to note that at the particular runway used for this test, there is a

"hump" in the runway at about the midpoint. That is to say that the runway

elevation is higher in the middle than on either end. With the pilot closing on

radar altitude, the maneuver proceeds nicely until that point, at which time a

power adjustment is required as the runway "falls away" from the airplane. This

"feature" in the local topography provi ded a convenient increase in workload for

the pilot flying the task.

122
Comments on Use of Simulation

• Most Valuable for Pilot Familiarization and Practice of

Maneuvers

• Easy to Measure Pilot Performance

• Lack of Cues Makes PrecisionTasks More Demanding

- Depth Perception

- Visual Acuity/Scene Content

- Motion

• Lack of Urgency Allows Higher Pilot Gain

• PIO Results are Largely Inconclusive

At this point, a small diversion into the subject of the use of simulation is in

order. Boeing uses engineering simulation, with pilots in the loop, both fixed and

moving base for this kind of testing. As a result of this experience, these sessions

are seen as more valuable for pilot familiarization with the task than for

collecting data regarding APC tendencies of a particular configuration.

While it is easy to measure and feed back pilot/vehicle performance in the

simulation, there are a number of deficiencies as well. On-ground simulation is

simply not the same as flight. A number of pilot cues, which may or may not be

important for a given APC evaluation are lacking or of insufficient quality. In

addition, the pilot knows it is a simulation, and so there is a general lack of

urgency. Pilots have been seen to make control movements in simulation which

they simply would not do in flight with a large transport.

Based on this experience, PIO results from simulation alone are considered

largely inconclusive.

123
Simulation / Flight Performance

Simulation Flight

'T

'° • I

_° iI

I

.... 1

1

J

j

i

_°

.°

H°

One example is shown in thi s comparison. On the right is the in-flight result

from the straight fly-by maneuver shown previously. On the left is a time history

taken in a fixed base simulator. For whatever reason, the pilot is simply not able

to fly the required task in the simulator.

Use of simul ation can certainly flag the potential for untoward tendenci es, but

the effects of myriad cueing issues are yet unanswered. As a result, ground-based

simulation is not yet seen as a viable substitute for flight testing. However, it is

quite valuable in getting pilots familiar with the maneuvers involved and useful

as a tool to explore maneuver set up, etc.

124
Lateral S-Turns

I

I

L

Intended to Increase Workload I

I

I .

by Adding Axis ......,

/ I

- Fly ILS to 50 Feet " ,'

I

- Acquire as Rapidly as Possible

one Runway Edge Line ', '

- Acquire as Rapidly as Possible '

the Opposite Edge Line

i

- Repeat for Length of Runway [

- Maintain 50 +/- 10 Feet

- PNF Calls Radar Altitude

f

t,L , dq-a

In an attempt to increase the workload encountered on the fly-by maneuver, an

additional task was superimposed. The lateral S-Turn maneuver asks the pilot to

proceed as in the flyby, except once established at 50 feet, the pilot should, as

rapidly as possible acquire alternate runway edge lines and continue for the

length of the runway.

This is a very impressive maneuver for an airplane with a 200 foot wingspan at

50 feet above the runway.

125
Vertical S-Maneuvers

• Further Increases Urgency

- Fly ILS to 50 Feet and Capture 50 +/- 10 Feet

- Acquire as Rapidly as Possible 30 +/- 10 Feet

- Acquire as Rapidly as Possible 70 +/- 10 Feet

- Repeat for Length of Runway

- Maintain Centerline

- PNF Calls Radar Altitude

_-::;-..... r.... _-_ ............ __I_o,,,,

,oF°°,__:,°Fee, l

An additional increase in urgency was achieved when the pilots were asked to

perform a vertical S-maneuver. Again leveling at 50 feet, the pilot is asked to

rapidly and aggressively acquire 30 feet and 70 feet alternately. While this is a

single axis task, urgency is very high in a large airplane maneuvering vertically

close to the ground.

126
Offset Precision Landing

I

I

I

I

I

I

1

ix...

I

Fly ILS at

2 Dots Offset

2 Dots High

Correct at 250 AGL

Land On Centerline

N_n Touchdown Zone

The offset precision landing is a maneuver used by most testing organizations to

investigate PIO tendencies, and Boeing has used it as well. The familiar set-up

for this maneuver is to align on the drainage ditch beside the runway at Buffalo,

NY, as used by Veridian/Calspan. Most airports do not have this convenient

landmark, however, so Boeing has adopted a multi-axi s task which involves

flying the ILS intentionally offset. The offset chosen is 2 dots laterally and 2

dots high. At 250 AGL, the pilot is asked to correct to the centerline and land in

the touchdown zone. This is a very challenging maneuver at low altitude.

127
Flyby / Landing Evaluation

Summary

• Combines Acquisition with Tracking

• Very Demanding Piloting Tasks

• Urgency is High Near the Ground

• Performance is Measurable / Readable

• Regarded by Some as High Risk

For the low altitude tasks, Boeing has chosen maneuvers which combine

acquisition with tight tracking in very demanding tasks. Being close to the

ground increases the pilot's urgency and thus pilot gain. Because the target (the

runway) is fixed in space, it is relatively easy to measure quantitative

pilot/vehicle performance.

A consideration worthy of note is the proximity to the ground with a very large

airplane is regarded (properly) by some as high ri sk. The risk of encountering

undesirable characteri stics in such a situation must always b e weighed in the test

planning process.

128
Other Maneuvers in the Toolbox

Flight Director Tracking

- Sum-of-Sines

- Steps-and-Ramps

- Log Frequency Sweeps

- Added Discrete Disturbances

Bank Angle Captures

Heading Angle Captures

Lateral Pilot Handoff

Full Rudder Sideslip in Ground Effect

Constant Track Rudder Step

While the "generic" maneuver set is defined as above, a number of other maneuvers

have been used for specialized appli cations.

Flight Director Iracking has been used in some cases, with a number of different input

functions. In all cases, the pilot is shown only the error between commanded attitude

and actual attitude, forcing a compensatory tracking scheme. Log frequency sweeps

provided both insight and broad frequency coverage for future analysis. The ability to

insert discrete disturbances into the flight director signal also provided additional insight.

Bank angle and headi ng angle captures are standard evaluation m aneuv ers. The lateral

pilot handoff i nvol yes one pi lot initi atin g a roll i ng maneuver, relinquis hi ng comm and of

the airplane to the other pilot while at the same time calling out a bank angle to capture.

This is essential ly a bank angle capture initiated from a non-zero roll rate.

Full rudder sideslips in ground effect are an attempt to investigate a landing de-crab

maneuver in much the same way that the fly-by allowed investigation of the landing

flare.

The constant track rudder step is an up-and-away maneuver in which the pilot inserts a

rudder step and flys track (on the nav display) with wheel. This maneuver turned out to

be very difficult to fly. While it is essentially a transition from crab to slip as in a

crosswind landing, it proved unnatural to perform up and away on instruments.

129
Flight Test Evaluation Summary

• Boeing has Extensive Experience Flight Testing for PIO

- Several Hundred Hours of Testing

- Six Different Models

- Large Number of Manuevers / Techniques

• No Single Maneuver / Technique has Proven to be

Effective for Exposing PIO Tendencies

• Most Effective Testing Strategy Appears to be Careful

Diligence During Normal Test Flying

• Prudent Handling Qualities Design Appears to be Effective

for Prevention

• Evaluation Process Continues to Evolve

Through several hundred hours of flight testing to evaluate PIO tendenci es over a

large numb er of airpl anem odels an d in volv ing a large number of speci fi c

maneuvers, no single maneuver or technique has proven to be effective for

exposing potential PIO tendencies. The conclusion from this is that the most

effective design strategy appears to be prudent attention to fundamental handling

quali ties design while the most effective testing strategy appears to be careful

diligence during normal test flying. The testing which is done for development

and certification of a transport airplane provides significant opportunities to be at

remote comers of the flight envelope and investigate airplane characteristics.

Even so, the evaluation process continues to evolve and more new information is

learned with each additional test program.

130
Fw

Response Linearity

Fw 8w

_W

,,

J

as

Moving from generic testing to identi fying challenges for future work, this chart

depicts a number of steps between the pilot's application of force to an inceptor

and the airplane response.

In the upper left is a (crude) depiction of a column/yoke. As the pilot applies a

force (Fw) to the wheel, the wheel would be expected to move. Moreover, as the

sketch below it shows, it is normally assumed that there is some linear

relationship between appl led force and wheel deflection Ogw).

For mechanical or di splac ement command systems, that displ acem ent of the

wheel should result in a corresponding di splacement of an aerodynamic surface

(_is), as depicted in the center sketch. Again, it is typically assumed that there is

a linear relationship between controller displacement and surface displacement,

as in the sketch in the upper right corner.

Finally, a surface displacement (_Ss) is expected to result in an acceleration of the

airplane, in this case, a roll acceleration (0"). In most cases there is a goal to

achieve a linear relationship between these two as well, as shown in the lower

right sketch.

These assumptions of lin earity form the basi s for the use of frequency domain

analysis to study airplane dynamics and PIO.

131
Fw

Real

Fw _w

World

.A

/ i

_w

(Non)Linearity

_S

_W

/

J

,/

J 5s

Unfortunately, the real world does not always conform to these assumptions.

In the presence of system friction, the control force to controller displacement

relationship exhibits discontinuities and hysteresis. (lower left).

Modem transport airplanes typically use a combination of ail eron and spoil er

surfaces for roll control, each of which may be scheduled on different deflection

curves, have different rate capabilities, etc. (upper right)

Finally, though a linear roll rate capability is desired, it is rarely achieved in

practice.

Each of these sources of nonlinearity causes difficulty in application of the

typical analysis methods for PIO which are found in the literature. To focus on

the need for methods to accommodate these characteristics, each is discussed in

detail in what follows.

132
Controller Characteristics

Fw

Breakout

Force

,j. ...-

Friction___

/f!

t i

JJJ i

t.1...I..." J

/ J

Control

Centering

_w

Starting at the pilot's fingertips, while most agree that linear force/displacement

characteristics are des irable, al 1 control sy stems h ave fricti on. In particular, large

transport aircraft with m echanic al control sy stems can have friction level s which

are not trivial.

One thing that friction brings is hysteresis. In order to achieve some degree of

control centering,, a breakout force is typically added. This breakout essentially

offsets the force/displ acem ent curves around zero, allowing the wheel to return to

the center positi on when no force is applied.

133
Friction Generates Gradient

Ambiguity Around Detent

Fw I / G°w._/" ._

Force -- l_j/

J_

_w

j..._

.I j-/

A r // I_ Detent Gradient

Friction___ k./j/__ /

The presence of this breakout produces a force/displacement disconti nui ty. The

presence of a slope change can have detrimental effects on pilot p redictabi lity.

The pilot loses his sense of how much force to apply to get a desired

displacement. Moreover, the slope discontinuity is right in the center of the

control operating range, where the pilot works the most. This can make small

displacements, e.g. those required for tight tracking around neutral wheel,

difficult for the pilot.

134
Gradient Ambiguity Away From

Detent is Function of Amplitude

JJ J fl tJ_

Fw_._21& j/-

.z/ /

_W

Away from the detent, the presence of friction and the associated hysteresis

causes a similar gradient ambiguity. Moreover, the degree of ambiguity is a

function of the size of the input for a given friction level.

This is significant for example in a decrab maneuver for a crosswind landing.

The gradient of the force required to move the wheel a given amount in each

direction around a (non-zero) trim point depends on how big the input needs to

be.

Again, predictabil ity from the pilot's point of view is compromised.

135
Dynamic Inertial Effects on

Controller Characteristics

.°,,,°, °,,,,,,,

The static force/displ acem ent characteristics of the controller are only part of the

story. Since the control system itself has mass (and large transports can exhibit

significant mass characteristics), the force/displacement characteristics vary as a

function of the frequency or speed at which the control is moved.

What is shown is force vs displacement at near zero frequency and another sweep

at signi fi cantly high er frequency. It is clear that the two curves are si gnifi cantly

different. The center detent is not even evident in the high frequency case, the

slope of the return (long lower path going from right to left) at hi gh frequency is

not similar to the near zero frequency case, and there are some n on-li near

characteristics near the ends of the travel.

136
Dynamic Inertial Effects Depend

Also on Path (History)

e_

Now, the high frequency sweep on the previous chart was taken from the middle

of a log frequency sweep. Had a single high frequency sweep been undertaken

from a standing start, the force/displacement curve would have looked different

yet. All of this is because the control system itself has mass and inertia.

137
Dynamic Inertial Effects on

Controller Characteristics

/

The end result is again a question of predictability. At any given time in the

flying of an airplane, the pilot needs to have some idea of how much force to

apply to the controller to get to move to where he wants it to go. These dynamic

characteristics cloud the issue and contribute to ambiguity.

138
Control Activity on Final

Approach

What this has to do with real flying of airplanes is shown here. This is a time

history of wheel position for a normal approach to landing. Wind was light,

turbulence was not a factor.

What is unique about this is the pulse-like character of the wheel inputs. At the

left hand side note the quick pulse as the wheel moves more than 15 degrees,

then is taken back to zero in about a half second. This is followed by an equal

pulse in the other direction. After a period of quiescence, the sequence is

repeated at roughly twice the amplitude, still with very short duration.

139
Pilot / Controller Interaction

iI

i i i I ,i i i I I .o_ ,_ , ,o +,, , ,s _ ,. ,.i . ,4

wi.o- |lil_iI

Just why this is happening can b e further understood by exami ning the

corresponding pi lot force inputs.

Note that between the first and second position doublets, where the wheel is

approximately zero, the force is not. In fact the pilot fried to move the wheel.

There is a brief 5 pound input in which the wheel did not move. This is followed

by a larger, nearly 10 pound input which generated the larger wheel deft ection

(upward on this plot) which the pilot immediately removed, and corrected in the

other direction.

In this case, the wheel feels "sticky" to the pilot and small, smooth inputs are

difficult. This degrades precision of control.

140
Effective Controller

Characteristics

,z

.q

j i .... "

-.. -z _ -Ib • .o Je ze

_..,,.Z P.*,,...

A phase-plane representation of the same sequence is overlaid on the near-zero

frequency force/displ acem ent plot for the same configurati on. This illustrates the

lack of predictabil ity which is generated by i nertial ch aracteristics of the control

system itself.

The result is that at any point in this dynamic maneuver, the pilot is unable to

predict how much force to apply to generate what wheel position.

These kinds of controll er effects are not adequately dealt with in the literature,

and represent an area which is ripe for investigation.

141
Determine "Best" Controller

Characteristics Set

• Given Minimum:

- System Inertial Characteristics

- System Damping

- System Friction

• With Constraints on Maximum:

- Force at Stop

- Power to Drive System (Pilot Qualitative Input)

• Find Desirable Combinations of Breakout,

Gradient, and Damping

These were dealt with at Boeing in the following way.

It is understood that the control system has a minimum inertia, damping, and

friction. Any modifications cannot change those, although additions to each

would be possible.

In addition, there are constraints on maximum force at the wheel stop

(regulatory) and on the power to drive the system (e.g. if friction or damping get

too hi gh, pilots will be easily fatigued by sim ply moving the wheel around).

The challenge was to find desirable combinations of these parameters to improve

the pilots ability to make smooth, predictable control inputs.

142
Human Centered Design

The Experiment

D_

K

An experiment was designed for a high fidelity simulation in which the control

loader characteristics could be changed to reflect the changes in the parameters.

This is a time history of the wheel deflections commanded in the study. The

pil ots were asked, to position the wheel according to this scheme.

This did not involve "flying" an airplane model at this point. It was simply a

one-dim ensi onal task to see if some combinations of friction, damping, and

inertia were better than others for the pilots' ability to precisely position the

wheel.

In looking at some results, the time period just after the full left wheel input will

be exami ned.

143
Human Centered Design

Some Results

)411

! _"_, !

:!, _.

!;l

• ,-,:

i_"

!-

i

1J

Some sample results are given here. In the time history plots, wheel position is

on the top, wheel force is on the bottom.

For the configuration on the left, it is clear that the pilot was able to achieve the

desired wheel positions accurately and quickly with little overshoot. Good

dampi ng is seen on the lower force trace, wherein the pilot used a small but well

damped oscillatory force input in order to get a good square shaped response.

For the configuration on the right, it is just as clear that the pilot is having

difficulty achi eving the desired wheel positions. The force oscillatory at the

comer points is not as well damped as before, and larger in magnitude.

144
Application of Results

• "Best" Configurations (and one "Bad" one)

Flown in Simulation for Pilot Opinion

• Best of Those Configurations Flown in

Flight Test

...Results Indicate Improved Pilot Opinion,

Improved Precision (Pilot Performance),

and Less Structural Excitation

With the results from the single axis wheel positioning task, the "best"

configurations were flown along with an airplane model, still in simulation,

asking the pilot to perform operational tasks. This was also done with one

configuration deem ed "bad" by the single axis task, just to in sure that the first

results were not mi sleading.

The best combinations of friction, damping, and inertia from simul ation were

fl own in flight test (airplane systems were modified to match the characteristics

determined in simulation).

The results of the flight testing indicated that pilots did indeed both prefer the

new feel configuration and found that it afforded them a higher level of precision

in their maneuver performance. An unexpected benefit was the realization that

with the new configuration maneuvers could be flown with less structural

excitation.

145
System Response Characteristics

_S

_w

As was mentioned earlier, on modem jet Iransport aircraft, the roll control

surfaces are often schedul ed separately as a function of controll er deft ection.

Ailerons and spoilers are often actuated on different schedules and with different

rate capabi lity actuators.

146
Effect of Frequency on System

Performance

/,;-'."

_..o.i p..,.i..

The presence of rate limits in any element of the system generates ambiguity

with respect to surface position which is a function of the frequency of the

controller motion.

Shown here is controller position vs surface position. For the near-zero

frequency case, the relationship is indeed close to linear. However, at larger

frequencies, particularly past that required to saturate actuator rate limits, the

relationship becomes more ambiguous.

To the pilot, this means that at any point in time, the surface position may not

correspond to the controller position.

147
System Response Linearity

Phase Delay is Amplitude Dependent

-30% -50%

Deft. Deft.

[] Normalized Phase

Delay Parameter

For cyclic motion of the controller, the rate limits are reached at different

frequencies for different amplitudes of motion. This will show up as a nonconstant phase delay parameter as a functi on of controller deft ection.

Shown here are results of frequency sweeps done at three different amplitudes,

indicating that at larger deflections, the apparent phase delay can become

significantly larger than at lower deflections. This can come as a surprise to the

pilot who had predi ctabl e characteristics with smal ler deft ections.

148
Aerodynamic Response Linearity

Generates Gradient Ambiguity

t

8s

The final element in the nonlinear control response story is the aerodynamic

response to surface deflection. While it is desirable to achieve a linear response

to surface deft ection, such is simply not always the case.

For the same reasons that the control force characteristics produce ambiguity,

discontinui ties in aerodynamic response do as well. For example, consider a

pilot holding a sideslip requiring a surface deflection between the two yel low

points. Correction for gusts which may force a deflection which crosses one or

both points, will result in the pi lot geting less response than was commanded

based on the first seen gradient. This lack of predictability can result in loss of

precision and frustration on the part of the pilot.

149
The Result Is Really Difficult to

Analyze

• Modem Airplanes Have Many Nonlinear

Elements

• Pilots are Quite Adaptable Controllers

• Current Theory is Inadequate for these

Cases

The end result of all of these nonlinear elements is of course that the real airplane

is real ly difficult to analyze with current methods.

Complicating the situation is the fact that pilots, and in particular test pilots, are

remarkably adaptable controllers. They may compensate for these el ements

without being aware that they are, and they may not be able to communicate to

the engineer the full consequences of the situation.

Finally, the state of the art in analytical techniques is not felt to be to the point at

which these elements can be addressed adequately, and in particular with regard

to PIO tendencies.

150
Pilot / Management Perceptions

There's a Fine Line Between:

Looking for a PIO_

Proving That There's Not

J One There

Ultimately, the pilot is on the spot to pass judgment on PIO tendencies.

Often, the pilot (and sometimes managers who listen to them) will b eli eve that

the engineer wants the pilot to induce a PIO. In fact, the engineer usually wants

to demonstrate that the pilot will not induce a PIO. The difference between these

two situations is often very fine.

In any case, encountering such an event is usually seen as an honest-to-goodness

out of control situation, which is generally considered not a good thing. Arriving

at an agreed upon set of conditions which will both adequately explore the

pil or/vehicle combination and retain adequate safety margin s is a very important

step in the process.

151
The Pilot is Part of the Equation

• Pilot "Gain" is Important in Closed Loop

Performance and Stability

• Pilot "Gain" is not Easily Controlled

• Standardized Evaluation Tasks will Require

a Consistent Level of Pilot Agressiveness

A very important part of the pilot/vehi cle combination is of course the p il ot

himself. An important part of the stability of the combi nation is the pilot "gain".

Unfortunately, most pilots don't change their gain at will. A few can increase

their gain when asked, but it is rare that a pilot, once in a "high gain" situation

can choose to reduce it.

If a standardized evaluation is to take place, there must be a way to normalize

pilot aggressiveness across pilots and across individual evaluations. This is

essenti al precisely becaus e of the extreme dependence of the result (PIO or no

PIO) on pilot gain.

152
Tecba..ques to Boost

Aggressiveness

• Maneuver Performance Requirements

- Extreme Precision in Performance

- Mandatory Control Positions (on stops)

• Urgent Flight Situation

- Close to the Ground

- Close to Another Airplane

• Consistency is Difficult to Achieve

Given what was said above about aggressiveness, it should be noted that there are

known ways of increasing an individual pilot's gain in a given situation. These

include maneuver performance control and control of the urgency of the flight

situation.

What remains uncertain, though is a way to achi eve consistency. Without that,

consistent evaluations will be difficult to achieve.

153
Validation Dilemma

• Evaluations must:

- Identify PIO Prone Configurations

- Pass Configurations Which are Not PIO Prone

- Give Consistent Results Across Pilot

Populations

- Be available without undue cost/schedule

impact

• JAA/FAA/Industry are Working Together

What can be said about techniques for validating that a configuration is free of

PIO tendencies is what an evaluation criterion must do.

Accurate identification of PIO prone configurations is obviously an important

characteristic of any evaluation technique.

Equally important is the ability to pass configurations which are not PIO prone.

False positives can result in wasted tim e and energy in identifyi ng unnecessary

solutions.

Any proposed evaluation technique must give consistent results across pilot

populations so that the results do not depend on which pilot does the evaluation.

Finally, any evaluation technique should be available without undue cost or

schedule im pact.

The dilemma is of course that there is no evidence that an evaluation metric is

available which meets these criteria.

The good news is that the world's regulatory authorities for transport aircraft are

actively working together to monitor the situation and act if appropriate.

154
Summary

• Boeing's Experience in Testing for PIO is Extensive

- Generic Testing Program is in Place

- Database is Being Built / Lessons are Recorded

- Toolbox is Growing

- Effective Validation Maneuvers are Elusive

• Many Analysis Details are Available for Consideration

• Most Effective Prevention Strategy is Prudent Handling

Qualities Design Practice

• Pilots Are a Key Ingredient: They Must be Involved

• Most Effective Testing Stragegy Appears to be Careful

Diligence in Normal Test Flying

• The Process Continues to Evolve

155

PHANTOM WORKS

Stability, Control & Flying Quafities

The Effects on Flying Qualities and PIO of

Non-Linearities in Control Systems

Edmund Field

Boeing, Long Beach

PIO Workshop

NASA Dryden

April 6-8, 1999

*a', //// Ou{_,ut

A _A-a

Hystefes*s Descrip_ on

h t J_J_

• . /-C)ulDLII

/

\

Irlpul _,ld Oulput _n'te Histories

Factors that cause Category I PIOs have received much attention over many

years, resulting in the development of many PIO prediction criteria.

More recently attention has turned to Category II PIOs, those that include nonlinear effects such as rate limiting. Other sources of non-linearity also exist in

an aircraft's control system, however, these have received less attention.

This presentation discusses some recent experience with non-linear elements

in control systems, and their implications for flying qualities and PIO

susceptibility.

157
_ PHANTOM WORKS

The Effects on Flying Qualities & PlO of Non-LJnearities in Control Systems

Sf_lbJlAry, Control & Flytr_ Q_Jaf_ttes

Background

Most Flying Qualities and PIO criteria assume linear models

for all elements in the total control / aircraft system

Fes

-- .FeelSystem _es Mecha Linkagesnica, _ Actuator ___

Aircraft

0

Y

n z

_ JOfPNO

Most flying qualities and PIO prediction criteria assume linear models for all

elements in the total control / aircraft system. That includes linear models of

the feel system, the mechanical linkages, the actuators and the aircraft

dyn ami cs.

Category I PIO criteria concern only linear causes of PIO.

Category II PlO assume non-linearities due to rate limiting only, all other

elements in the total control / aircraft system are assumed linear.

While this may be reasonable for a first approximation, in reality all these

elements include some non-linearities. The total contribution of all these nonlinearities may become appreciable and so have important implications for an

aircraft's flying qualities and PIO susceptibility.

For example, hysteresis in the feel system is a well known phenomenon, and

yet its effect on an aircraft's flying qualities are neglected when performing

linear analyses. To some extent its effects can be neglected if the analyses use

control inceptor position (as opposed to force) as the input. However, the

effects of the hysteresis should be taken into account elsewhere. Current

criteria for this are lacking.

158
<_ PHANTOM WORKS

$eab_l Comrol & FlyJr_ Qua#t_s

The Effects on Flying Ousflties & PIO of Non-Linearitles in Control Systems

Analysis of Pitch Frequency Sweeps Identified Phase Loss at

all Frequencies

This phase loss may have been caused by non-linearities in the control system

-50 ...... : ( -- Linear Model

J

-60 ............. _.......... Non-Linear Model i

-80

-90

01 1 10

'8°L I

01 1 10

Frequency

When analyzing data obtained from pilot generated pitch axis frequency

sweeps a phase loss was identified at all frequencies in the Bodes of stick force

to aircraft response. It was suggested by Mr. Dave Mitchell that this phase loss

may have been caused by non-linearities in the control system, specifically

hysteresis.

159
<_ PHANTOM WORKS

The Effects on Flying Qualities & PIO of Non.Linearities in Con trol Systems

Stabile, Control & Ffy_ng Q_JaMbs

Categories of Non-Linearities

InpuVOutput Sirnl_e _

Relationship

Phase Amjle Zero Non-Zero

Arrplitude Yes Yes

_?

Frequency No No Yes

r:xan_es: Fna_an Hyslaeds Backlashwith

_d Toggle Coularrb

!SakralJon Bemanlary Fri_on

Baddash

There are several categories of non-linearity that may be present in an

aircraft's control system These may be represented by either simple or

complex describing functions ].

Simple non-linearities exhibit gain attenuation, but no phase attenuation. The

gain attenuation is independent of the frequency of the input, but dependent

upon the magnitude of the input amplitude. Examples include friction,

threshold and saturation.

Complex non-linearities exhibit both gain and phase attenuation. The

magnitude of the gain attenuation is dependent upon the magnitude of the

input amplitude, and may or may not be dependent upon the frequency of the

input. Examples of frequency independent complex non-linearities include

hysteresis, toggle and elementary backlash. Frequency dependent nonlineari ties include backlash with Coulomb friction.

Various of these non-li nearities may be present in an aircraft's control system.

When added together, from the pilot applying a force to the control inceptor to

the aircraft responding, there may be appreciable gain and phase attenuation at

all frequencies.

i Graham, Dunstan, and McRuer, Duane, "Analysis of Nonlinear Control Systems", John

Wiley and Sons, 1961

160
_ PHANTOM WORKS

SfabJIlly, Control • Flyk V Quai

The Effects on Flying Qualities & PIO of Non-Linearities in Control Systems

Hysteresis - An Example Control System Non-Linearity

+A L-Inpul

Hysteresis Description

+A-a

/

+a !

-a

-A+a

-A

. L.

Output

time

Input and Output Time Histories

Hysteresis is a well known non-linearity which is present in aircraft feel

systems. The effects of hysteresis will be discussed as a representative

exam ple of control system non-li nearities.

Hysteresis is a complex non-linearity which produces gain and phase

attenuation independent of the frequency of the input.

In the following discussion the characteristics of hysteresis will be described

by the magnitude of the non-li neari ty 'a' and the magnitude of the input signal

The effect of the non-linearity in the time domain is evident in the figure. The

magnitude of the output is limited to 'A-a', and the output is lagged behind the

input, as well as the shape being modified.

The magnitude limiting causes the gain attenuation and the lag provides the

phase attenuation that is evident in the Bode plots.

161
<_ PHANTOM WORKS

Srablll_,, Control & Flying Qualities

Sinusoidal Describing Function for Hysteresis

The Effects on Flying Qualities & PIO of Non-Linesrities in Control Systems

.s_° i t' ! i . ii

_,2oi i __i "_

]

T f i

J r L

° :!' -lo ' - i

.:m . i

_.o ! ....., i

< i l

_ , ! '

i i t,,

ir I ! _

The sinusoidal describing function for hysteresis is shown graphically. The

magnitude of the gain and phase attenuation provided by the hysteresis is

simply a function of the ratio of the magnitudes of the non-linearity to the

input, 'a/A'.

When 'a/A' is zero (i.e. zero deadband) there is no gain or phase attenuation.

As 'a/A' increases both gain and phase loss increase as the effect of part of the

applied force is now lost in the deadband zone (-a to +a). As 'a/A' increases

towards 1 (all applied force is in the deadband region) the gain and phase

attenuation approaches infinity, there is no output to the corresponding input.

162
_ PHANTOM WORKS

Stablll_/, Costtrol & Flying OualttMs

The Effects on Flying Qualities & PIO of Non-Llnearities in Con trol Systems

Time Histories from Typical Piloted Sweep

Input Magnitude (A) Increases as Frequency Increases

,_ ...._...!..., .!.--!..-_ ..:.. ! !.+_....!+ _- !.......

i -4

T;mc (=E¢)

4 ............ 1

_,I+++7+7i°++! ! +++! ! +I

io _ Jo d¢ _o Io 7o io io ioo .o _0 1_ _ i1o

Tim. (+_)

4

+ : : : : i *+ + i i :

,+..... ++++H+ ............ +++ ...... i+?+i ++

Although hysteresis is a frequency independent non-lineafity, the attenuation it

introduces may v ary wi th frequency i ndi rectly.

The figure shows time histories taken from a typical piloted frequency sweep.

It can be seen from the figure that as the frequency of the pilot inputs increases

the magnitude of the inputs ('A') also changes. Generally, as the frequency

increases so does the magnitude, although this is not universally true.

The implications for the analysis of frequency sweep data is that the

attenuation introduced by any non-linearities may be affected by the

frequency/magnitude rel ationsh ip of the i nput.

163
<_ PHANTOM WORKS

Stability, Con_'ol S Flyin_ Qualities

The Effects on Flying Qualities & PIO of Non-Linearities in Control Systems

Gain & Phase Attenuation Relationship to Deadband Magnitude

/

5 io 15

nout Conlr_ Column Force i1_)

o .r'" -- -- I

c

.14

16

2o 25

o-

.lo

15,-2o

.5_-

o-_0-

f" . ., -, t/

Input Conlr_ Column Force {_)

2O 25

12 Ib D,eadband

10 Ib Deadband

8 Ib Deadband

6 Ib Deadband

4 Ib Deadband

2 Ib Deadband

0 Ib Deadband

The gain and phase attenuation provided by hysteresis is a function of the

magnitudes of the non-linearity 'a' and the input sinusoid 'A'. During a

frequency sweep, such as that shown on the previous slide, 'a' remains

constant, but 'A' varies, possibly with frequency. The figures show the

variation in gain and phase attenuation with input magnitude 'A' for 7

different values of non-linearity 'a'. Also included are lines of constant 'a/A',

taken from the sl ide before the previous.

For a constant deadband 'a', as 'A' increases 'a/A' will reduce. This can be

seen by following a line of constant deadband, for instance the solid bold line

for a deadband of 8 lb (a = 4 lb either side of trim, to give a total deadband of 8

lb). For low force inputs 'a/A' is high, about 0.9 at 4.5 lb. As the magnitude of

the inputs increase 'a/A' reduces, so that at 6 lb input 'a/A' is 0.7, at 8 lb 'a/A'

is 0.5 and at 13 lb 'a/A' is 0.3. As the force increases and 'a/A' decreases the

curves of constant deadband flatten. The change in gain and phase attenuation

with increasing applied force becomes minimal. Physically, this is because the

effect of the deadband becomes reduced as the available applied force 'A-a'

becomes much larger than 'a'.

The implications for piloted frequency sweep generated data are that the gain

and phase attenuation introduced by the non-linearities will be dependent upon

the magnitudes of the input, and to some extent will vary with frequency. This

make s the prediction of the effects of the non-linearities more difficult.

164
PHANTOM WORKS

Stmblllly. Con_n3; & Flying Oua#_s

The Effects on Flying Qualities & PIO of Non.Linearities in Control Systems

Implications for Flying Qualities and PIO Susceptibility

• The phase and gain attenuation introduced by non-linearities in the

control system will have implications for the flying qualities and PIO

susceptibility of the aircratt

• The gain and phase attenuation will be greatest for small control

inputs, such as during fine tracking tasks

• Non-linearities in aircraft control systems should be minimized to

reduce these effects

• Caution must be taken when applying flying qualities analyses

The phase and gain attenuation introduced by non-linearities in the control

system will have implications for the flying qualities and PIO susceptibifity of

the aircraft.

The greatest attenuation will be observed when making small control inputs,

such as during fine tracking tasks. Susceptibility to PIO will be greatest for

these tasks.

Where possible, the non-linearities in aircraft control systems should be

minimi zed to reduce the attenuation effects they introduce.

When performing flying qualities analyze it is important to appreciate the

effects that control systems non-linearities have on an aircraft's flying qualities

and PIO susceptibility. Linear analyses that exclude these non-linearities are

prone to error, and are likely to predict better flying qualities and lower PIO

susceptibility than the real aircraft will exhibit.

165
<_ PHANTOM WORKS The Effects on Flying Qualities & PlO of Non.Linearities in Control Systems

$1abil_y, Control & FJyJng Qualtt_s

Implications for Flyi ng Qualities Analyses

Aircraft Models:

• Usually linear models are used. They do not include phase

attenuation characteristics of non-linearities

Flight Data:

• Complete non-linear aircraft. Data does include phase attenuation

characteristics of non-linearities

• The effects of the non-linearities dependent upon the magnitude of

the control inputs

Inceptor Force or Position?:

• Control inceptor force or position can be used as input. Using

position avoids the effect of the inceptor hysteresis, a major

contributor to the phase attenuation

• Elements between the feel system and actuator will be present in

both force and position analyses { .,.,.,,,,,,,

Control system non-linearities introduce several implications for performing

flying qualities analyses. It is important that appropriate analyses are

performed and that criteri a are appli ed consi stently.

When analyzing aircraft models usually only the linear dynamics are

considered, and the non-linearities are neglected. Data obtained in-flight

represent the total non-linear aircraft. Care must be taken when comparing

results from analyses of the linear model and flight derived data. Additionally,

data obtained in-flight will be dependent upon the magnitude of the input.

The choice of whether to use stick force or stick position as the input for such

analyses will affect the results, since the feel system includes non-linear

effects such as hysteresis. Using stick position will limit the included nonlinearities.

The implications of analyzing data from the non-linear model (or flight

derived data) will be demonstrated against two popular flying qualities

analyses:

• Low Order Equivalent Systems

• Bandwidth Criterion

166
<_ PHANTOM WORKS

The Effects on Flying Qualities & PIO of Non-Linearitie$ in Control Systems

Sfabillty, Control& Fty_g QuaUI_s

Low Order Equivalent Systems (LOES)

To achieve a good match the LOES dynamics may be altered to account for the

phase loss. In the Pitch axis, particularly _ph, {sp, Te and perhaps mph, rasp. 1/Te2

-50

-60 ....... [-- L,.... Mode, l

..... -- Non-Linear Model i

-80

-90

01 1 10

180 /

90 .....

T_r_

g_ ---- T. t

-180 _ j !

01 1 10

Frequency

For a constant gain attenuation at all frequencies the only impact on the LOES

fit will be a lower gain factor. If the gain attenuation is not constant across all

frequencies then the poles and zeros may be affected, possibly resulting in

changes to the equivalent short period frequency and damping. Any phase

attenuation, regardless of whether frequency dependent or independent, will

result in different LOES matches between the linear and non-I inear models.

A constant phase loss across all frequencies will likely be matched by an

increase in the equivalent damping ratios of the oscillatory modes ((sp and (ph),

spreading the phase reduction across a wider (and so lower) frequency range.

If this alone is unable to provide sufficient phase loss it may also be necessary

to reduce the equivalent frequency of the oscillatory modes (wsp and _ph)-

Additionally the numerator term l/Toe may also move, partly to offset the

movement of the poles. The equivalent time delay term, To, will be adjusted to

account for any high frequency offset that is either residual from or caused by

the movement of the poles and zeros. Note also that To will also be affected if

there is any frequency dependent gain attenuation that causes movement of the

poles and zeros.

tOspand 1/To2, are both factors in CAP. A PIO prediction criterion based upon

CAP and T o has been proposed. Clearly, any inaccuracies in the prediction of

these parameters will affect the prediction of an aircraft's susceptibility to PIO.

The likely effect of hysteresis is to increase an aircraft's PIO susceptibility.

1(
_ PHANTOM WORKS

The Effects on Flying Qualities & PIO of Non-Linearities in Con trot Systems

StabllRy, Contr_ & Fly_ Qua#ties

Bandwidth Criterion

To account for the phase loss the Bandwidth frequencies (both attitude and flight

path) will be reduced. % may be affected, depending upon the type of nonlinearity.

-20- 1 _--- Linear Model I

-40 J -- Non-bnear Model

_=-so- i - i

-100

-t2o r

1 I 10

_9

..................... _( _)p,:,,

!

.2,o_ ; \ /_ / ,'o

01 Frequency _J.)l a 0 20_8o

As with LOES, a constant gain attenuation at all frequencies will not affect the

Bandwidth criterion parameters. Even if the gain attenuation is frequency

dependent it is unlikely to affect the Bandwidth criterion parameters since

most aircraft are phase Bandwidth limited, and whatever causes the gain

response to attenuate is likely to have a greater effect on the phase response.

Any downward shift of the phase response will have a direct effect on the

Bandwidth frequency, reducing it by 3rosw. Since rp is proportional to the

slope of the phase curve between 0918o and 2(ol8 o it will be affected slightly by

a downward shift in the phase response, as can be seen in the figure. However,

r e may be affected even more if the slope of the phase response is dramatically

different between the o918o and 2w18 o frequencies of the linear and non-linear

model s.

OaBw and r e are variables in a proposed PIO prediction criterion. Clearly their

accurate deft nition is important if the PIO prediction criterion is to be valid. As

with LOES, the omission of non-linearities from the analysis is likely to

predict the aircraft less PIO susceptible than it really is.

168
PHANTOM WORKS

StabtMy, Condor& Fh/tng Qualifies

The Effects on Flying Qualities & PIO of Non-Linearities in Control Systems

Conclusions

• Non-Linearities in control systems can introduce gain and

phase attenuation

• Depending upon the type of non-linearity, the attenuation

may be frequency and / or input magnitude dependent

• FQ analyses performed with and without the non-linearities

will yield different results

• This may account for inconsistent predictions from flying

qualities analyses of linear and non-linear models and

flight data, and when including and excluding the feel

system

_ PHANTOM WORKS

Stabihty, Control ¢ Fl_ng Quah_tes

The Effects on Flying Qualities & PIO of Non-Linearities in Control Systems

Recommendations

Non-Linearities in control systems must always be

considered when addressing an aircraft's flying qualities

This might be achieved through the development of a

criterion accounting for all non-linearities in a control

system. This metric might be additive to existing criteria

169

@

Mitigating the APC Threata work in progress

Ralph A'Harrah

APC Work_h op

DFRC

6-8 April 1999

My Perspective

• What I would do if I was responsible for

- Research

- Design & Development

- Flight Test

- Certification

- Airline Safety

- Accident Investigation

... relative to mitigating the APC threat

@

171
Mitigating the APC Threat- ','_%"AT

Cat. II APC Research

• Task Identification

- e.g., a large ("over driving") correction to an upset, followed

by cl osed-loop control to get back on original flight path

Subject Identification

- e.g., APC evaluation results from naive "line" pilots compared

with experienced test pilots

Vehicle Identification

- Variable stability aircraft, or ground based flight simulator, or

actual airc raft

continues

Mitigating the APC Threat - _:"_AT_,_, ,.

@

Cat. 11 APC Research , continued

• Design and demonstrate a control system

that is free from Cat. II APC characteristics

for a wide range of surface rate limits (e.g.,

from 1% to 100% of the maximum

achievable surface rate)

172
Mitigating the APC Threat- @AT

Design & Development @

• Incorporatefavorite PIO criteria into Mark Tischler's

Conduit* Program to address Cat. I

° Minimize the actuator energy metric (cost function) in

Conduit C(_C_trol Designer's U_..nified I_nt_erface)

- to reduce probability of "over driving" beyond rate limits, a

Cat. II condition

- to increase actuator life

• Utilize tactile control feedback I on primary controls to

warn of approach to rate and/or position limiting, with

active stops to preclude "over driving"

continues

_analogous to NRC's col lective limit cu_ing, AvWk, p.53, 22Feb99

Mitigating the APC Threat - @AT

@

De._ign _ Develc_pment, continued

• Backup tactile control feedback on primary

controls design with adaptive filtering 1.2to

compensate for time delay caused by "over

driving"

• Isolate pilot controlled surfaces and actuators from

non-pilot controlled surfaces and actuators

- Reduce erosion of pilot control response and

authority from non-piloted intrusion

IHanke, Dietrich, Phase compensation: a means of preventing APC caused by rate limiting,

Forschungbericht 98-15

ZRunqudqwist, Lars, Phase compensation of rate limiters in JAS-39 Grippen, AIAA Paper

96-3368

4

173
Mitigating the APC Threat- _:_,,,_'_ZXT,

Ground/Flight Test

• From ground calibration tests, detelmine the cockpit

controls to surface response time delay and hystersis

characteristics for inputs up to the maximum input

rate & deflection capability of the pilot

• If values exceed expectations/guidance

/specifications, evaluate options for improvement

• Alternately, evaluate on variable stability aircraft

while pelfomling off-set landing, large upset

correction, etc., Cat. 2 APC maneuvers to define

criticality of the pioble m

Note: The issue here is the consist ent abil ity of line pilot s to accommodate the chmRe in ti me

delay and hysteresis charactedstics that may be exparienoed as part of a"hai r raising"

expa-ience such as a large upset, oran eminent inflight

4

Mitigating the APC Threat- _:i_)AT

@

Continue APC exposure/training of certification pilots,

using a variable stability aircraft

Emphasize the determination of evaluation tasks for

Cat. II APC that are both safe and effective

Evaluate in flight APC Cat. I characteristics using

existing FAA APC testing bench mark tasks

Would not attempt Cat. II in-flight evaluation until

safe and effective test technique is identified

continues

174
Mitigating the APC Threat- C3AT

@

Certification, continued

• From ground calibration tests, detemaine the cockpit

controls to surface response time delay and hysteresis

characteristics for inputs up to the maximum input rate

& deflection capability of the pilot

continues

5

Mitigating the APC Threat - <_AT ':_:

@

Certi fi enti cm, continued

• If time delay or hysteresis values exceed

expectations/guidance/specifications, evaluate on

variable stability aircraft while performing off-set

landing, large upset correction, etc., Cat. 2 APC

maneuvers

Note: The issue here is the consistent ability of line

pilots to accommodate_ in time delay

and hysteresis characteristics that may be

experienced as part of a "hair raising" experience

175
Mitigating the APC Threat - ¢_i&,g

Airline Safety t_

• For the cockpit primary control inputs and the

resulting cont_l surface outputs, re coN at data rates

of 20 Hz orgieater on the QAR

• Initial APC Preculsor

- Monitor QAR data for the time lapse between reversal of

the cockpit control rate and the associated reversal of the

surface rate as APC prec ursor

• Flag occurrences with to> 100 msec.

• Flag & record values of towhen to >150 msec.

• Involve APC specialist for consistent flags, or

values oft D >150 msec.

continues

Mitigating the APC Threat- ,L_)AT

@

Airline gnfety

• Growth APC Precursor

- Utilize 20 Hz. or greater data rates on primary controls,

primary control surfaces, aircraft accelerations, and

warning, such as "stall" and "over-speed"

- Utilize QAR data to support Conduit as a monitor

° Flag occurrences violating Level 1 criteria.

• Flag & record values of t o when t o >150 msec., and

Level 2 criteria.

• Involve APC specialist for consistent flags, or

values of to > 150 msec

176
Mitigating the APC Threat -- :<'/A _":"' T

Accident Investigation

@

For the primary cockpit flight controls, the

ass ociated control surfaces, and aircraft

accelerations felt by the pilots, require that

crash recorders utilize data rates of 20 Hz or

greater

- when the flight crew is actively involved with

primary flight contlols

- when an emergency has been declaled

continues

Mitigating the APC Threat - <'_AT

@

Ace.ide, nt lnve_/ig:_ticm, continued

• In an investigation exhibiting significant crew

control activity, examine the time lapse between

cockpit control inputs, the associated control

surface responses, and accelerations (or other

response metrics, such as warnings) to which the

pilot may be responding

• If the time lapse exceeds 100-150 msec., include a

team of APC specialists as part of the investigative

team

177
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OMB No. 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

April 2001

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.PERFORMINGORGANIZATIONNAME(S)ANDADDRESStES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9.SPONSORING/MONITORING AGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

Conference Publication

5. FUNDINGNUMBERS

WU 529-55-24-E8-RR-00-000

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2407

10. SPONSORINGdMONITORING

AGENCY REPORT NUMBER

NASA/CP-2001-210389/

VOLI

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Unclassified--Unlimited

Subject Category 08

This report is available at http://www.dfrc.nasa.gov/DTRS/

12b. DISTRIBUTION CODE

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.

14.SUBJECTERMS

Flight control, Flight safety, Pilot-induced oscillation, Simulation of flight test

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