Development of longitudinal handling qualities criteria for large advanced supersonic aircraft

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NASA CR-137635

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DEVELOPMENT OF LONGITUDINAL HANDLING QUALITIES

CRITERIA FOR LARGE ADVANCED SUPERSONIC AIRCRAFT

FINAL REPORT

Robert W. Sudderth

Jeff G. Bohn

Martin A. Caniff

Gregory R. Bennett

N75-203_5

(NASA-C5-137635) DEVELOPMENT OF

LONGITUDINAL HANDLING QUALITIES CRITERIA FOR

LAPGE ADVANCED SUPERSONIC AIRCRAFT Final

Unclas

Report (Boeing Commercial Airplane Co.,

18212

G3/08

Seattle) 212 p HC $7.25 CSCL 01C

BOEING COHMERCIAL AIRPLANE COMPANY

FLIGHT CONTROLS TECHI4OI_OGY

L'

MARCH 19J i: '

Prepared Under Contract NA52-7966 by _...-

Boeing Commercial Airplane Company _k_/i,.- ....... >.'_,x5 _"

Seattle, Washington 98124

for

Ames Research Center, Moffett Field, California 94035

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

"== I

[ NASA CR-137635

4 I ,l,' jnd !_ jbT,Vh, 5 Repo, t [)at_,

MARCH 1975

DEVELOPIdENT OF LONGITUDINAL HANDLING QUALITIES

6 P,,r bL_rm,ng OtgJrllla|lO(I Coil{?

CRITERIA FOR LARGE ADVANCED SUPERSONIC AIRCRAFT

• _,,lh,)r ,_, }] Pl!,{{)llli,i_() Oll].{I}I/dl{()(l Repo,'_ N(J ROBERT W. SUDDERTH, JEFF G. BOHN,

MARTIN _. CANIFF, GREGORY R. BENNETT

141 W,)II_ [Jriil N(I

BOEING COMMERCIAL AIRPLANE COMPANY

P.O. BOX 3707

NAS2-7966

SEATTLE, WASHINGTON 98124

l _ I ,, p!" ,I R,_pO, t and Period Coveted

|

FINAL REPORT

',1 bl_on_.o_ n G A_)e'_._ Code

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON, D.C. 20546

I,_ ._Jl_ph!I_enl(]fv Nole_

Longitudinal handling qualities criteria in terms of airplane response

characteristics have been developed by The Boeing Company under contract

N_S2-7966. The criteria cover high speed cruise maneuvering, landing

approach and stall recovery. This report contains data substantiating

the study results which are reported in NASA CR-137635, "Development of

Longitudinal Handling Qualit,es Criteria for Large Advanced Supersonic

Aircraft.

I 8 [)lstr*blJtlO,l _tate,nerll

Handling qualities Supersonic cruise

Simulation SST

Pilot model FSAA

Minimum-safe Workload

.._11 re_r.ry__ _ La pr_a_

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TABLE OF CONTENTS

P_ q__,_ee

1.0 SUMMARY

2.0 INTRODUCTION

3.0 STUDY AREAS

,

4.0 TEST FACILITY

Ii

5.0 TEST AND ANALYSTS TECHNIQUE

5.1 PILOT RATING SCALES 12

5.2 PILOT MATH MODEL

5.3 WORKLOAD MEASUREMENTS

5.4 WIND MODEL 19

6.0 EXISTING CRITERIA COMPARISONS

7.0 STUDY RESULTS 32

7.] HIGH SPEED CRUISE MAHEUVERING

7.1.1 Pitch Attitude Display Sensitivity

7.].2 Pitch Rate _ivershoot Ratio 42

7.1.B Time-To-Peak Pitch Rate

7.1.4 Pitch Dampim] Constant

7.1.5 Column Force Gradient

7.2 LArlDING APPROACH (flORMAL OPERATION)

7.2.1 Pitch Rate Overshoot Ratio

7.2.2 Time-To-Peak Pitch Rate 71

7.2.3 Pitch Damping Constant

7.2.4 Column Force Gradient

a4

7.3 LANDING APPROACH (MIMIMUM-SAFE OPERATION)

7.3.1 Time-To-Double Pitch Attitude

7,3.2 Maneuver Margin

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7.3.3 Sudden Deqradation to Minimum Safe

7.3.4 Column Force Gradient

7.4 SIALL RECOVERY _!_TROL POWER

r,[

8.0 FOLLUW-ON STUDIES

9.0 CONCLUSION

_A'_.', _ -_-

9.1 HIGH SPEED _ MANEUVERI_IG

I06

9.2 LANDING APPROACH (_IORMAL OPERATION)

!99

9.3 LANDING APPROACII (MI_IMUM-SAFE OPERATIOII)

9.4 STALL RECOVERY COrITROL POWER

9.5 PILOT DESCRIBItIG FUNCTION STUDY

SYMBOLS AND ABBREVIATIONS

1:7

REFERENCES

i 1:!

APPENDIX A - SIMULATION FACILITY I;ESCR[PTIOI_

:26

B - PILOT MATH MO[]EL

C - WIHD AND TURBULENCE t,IODEL

:r,f

D - PILOT EXPERiE_CE

E - RUN SUMMARY

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L[SI OF FIGURLS

P

Paq.,.,__

TEST FACILITY

4-i Cockpit Configuration, First Simulation Study Period

4-2

Cockpit Configuration, Second Simulation Study Period

Electronic ALtitude Director Indicator"

4-3

TEST AND AIIALYSIS TECIINIQU[

5-] llandling Qualities Rating Scale

]_

5-2 Turbulence Effect Rating Scale

Pilot Workload Side Task Results with Turbulence

5-3

P

%

EXI_TING CRITEP, IA COMPARISONS

6-I SST Hiqh Speed Pitch Rate Response Criterion

6-2 SST lligh Speed Load Factor Response Criterion

6-3 SST Low Speed Pitch Rate Response Criterion

6-4 Shomber-Gertsen Longitudinal Handlinq Qualities Criteria

6-5 Cooper Rating Scale

9G

6-6 C* Hi(]h Speed Criterion

Jv

6-7 ;_inimum-Safe Calibration Time History

STUDY RESULTS

• j

7-i Pitch Rate Overshoot Ratio Responses,High Speed Cruise

3f_

7-2

Time-to-Peak Pitch Rate Responses, High Speed Cruise

")7

7-3 Pitch Damping Responses, High Speed Cruise

7-4 Load Factor Responses tu Pitch Rate Overshoot Ratio Tests,

High Speed Cruise

Load Factor Responses to Time-to-Peak Pitch Rate Tests,

7-5

Hiqh Speed Cruise

7-6 Load Factor Responses to Pitch []ampinq Tests, l!i,lh Speed

Cruise

7-7

Hiqh Speed Cruise Naneuverina Pilot Task

7-8 High Speed Cruise Evaluation of EADI Pitch Scale Sensitivity

7-9

High Speed Cruise Evaluation of Pitch Rate Overshoot Ratio

7-I0 Pitch Rate Overshoot Comparison with the SST High Speed

4_

Pitch Rate Response Criterion

7-11 Pitch Rate Overshoot Comparison with the SST _iigh Speed

Load Factor Response Criterion

7-12 MIL-F-8785B Criterion Comparison of lligh Speed Cruise

Pitch Rate Overshoot Patio

7-13 Hiqh Speed Cruise Comparison with C* Crite_iun (Baseline

Configuration)

7-14 High Speed Cruise Evaluation of Time-to-Peak Pitch Rate

7-15 Time-to-Peak Pitch Pate Comparison with the SST Hiqh

_.-_

Speed Pitch Rate Response Criterion

7-16 Time-to-Peak Pitch Rate Comparison with the SST Hiqn

!

CC

Speed Load Factor Response Criterion

I

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7-17

MLt.-F-_gUUB Criterion Comparison of High Speed Cruise

Time-to-Peak Pitch RaCe

7-18

High Speed Cruise Evaluation of longitudinal Damping

7-19

Pitch Damping Comparison with the SST High Speed

Pitch Rate Response Criterion

7-20

Pitch Damping Comparison with the SST High Speed Load

6C'

Factor Response Criterion

7-21

!AIL-F-8785B Criterion Comparison of lliqh Speed Cruise

Pitch Damping

7-22 63

Hiqh Speed Cruise Evaluation of Column Force Gradient

7-23 Landing Approach Pilot lask

7-24 Pitch Rate Responses, Landin,l Approach

&_

7-25

Landing ,_l_proach Eva_uation of Pi_ch Rate Oversho(,t :.:atic,

7-26

Pitch Rate Overshuot Comparison wikh th_ SST Low Sl_ei

7q

Pitch Rake Response Criterion

7-27 Pitcll Rate Overshoot RaLio Turbulence EvuluatioT_

7-28

Landing Apnroach EvalL_ation ot Time-to-Peak Pitt.h ;!at_ _

7-29

Time-to-Peak Pitch Rate Comparison witi_ the SST l_uw

7 t,

Speed Pitch Pate Response Criterion

-7 ,._

7_

Time-to-Peak Pitch Rate Turbulence Evaluation

7-31

Landing Approach [valuation of Pitch Damping

7-32 Comparison of Pilot Rating Prediction Technique

7-33 Pitch Damping Comparison with the SST Low Speed Pitc_ Rate

.,r_

Response Criterion

i

7-34[ Pitch Damping Turbulence Evaluation

7-35

Landing Approach Evaluation of Column Force Gradient

7-36 Column Force Gradient Turbulence [valuation

7-37

Minimum-Safe Landing Approach Time-to-Double Pitch

':,G

Attitude Evaluation

7-38 Effect of Turbulence on Landing _,pproach (Min-Safe)

Time-to-Double Pitch Attitude

7-39

Minimum Safe Landinq Approach Maneuvers Eva!uatlon

7-40

Landing Approach (Hin-Safe [valuation of Columr Crad_en{

;L

(T_ : 6.0 sac)

7-41 Lah_inq /_pproach {M_n-Safe _va]uaLion of Co]umn !,radiant

_C.

(T28 -- _.6 sac)

7-42

Longitudinal Stability-Stall Recovery Fvaluatio,]

: ",]

7-43

Stall Recovery Pilot Task

:52

7-44 Stall Recovery Control Power Evaluation

1_)3

7-C_5 Effect of Turbulence on St,_]] Recovery Control Power

CONCL US IOHS

9-I Modified SST iliqh Speed Pitc_ l_ate Response Criterion

!,--

9-2 Modified SST Low Speed Pitch Patc_ Resnons_ Criterion

9-3

Land_nq Approach (;.fin-Safe) Criterion

9-4

Stall Recovery Co_tT )l Power C, itPrion

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i- A ,_ , II I I Ill II -- _ ill

't

LlSl OF TABLLL

Paqe

qa,

7-! Hiqh Speed Cruise Maneuverinq Test Conditions

7-11 Landing Approach (Normal Operation) Test Conditions

7-111 landing Approach (Minimum-Safe Operation) Test

Conditions

7-1V Stall Recovery Control Power Test Conditions

P

P

D

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1.0 SUMMARY

A piloted simulation study has been made with the objective of advancing

the development of longitudinal handling Qualities criteria for farce suoersonic

cruise aircraft. This work was conducted on the Flight %insulator for Advanced

Aircraft (FSAA) located at HASA Ames Research Center. For this study the

simulator wos programmed with the math model representation ef the Boein_

2707-300PT Supersonic Transport as it existed at termination of the _ational

SST Program.

Areas of study included high speed cruise maneuverinq, landinq approach

for normal and minimum-safe operating conditions, and stall recovery control

power. The results of these evaluations were primarily based on oilot ratings.

Additional analysis capability was developed which consisted of a pilot model

analysis technique and pilot workload measurement techniques. The pilot model

results were obtai_,ed and utilized successfully for some of the conditions

evaluated in the l_ndinq approach (normal operation) study area. Pilot

workload was measured by two techniques; by a side task technique, and by a

computation of the physical work done by the pilot through the control column.

The side task technique was not successful. The physical measurement was

useful in analyzing landing approach conditions where major pilot rating

scatter existed.

The results of this study are a combination of new criteria and modifica-

tions of existing criteria. All pre-existing criteria utilized in the final

results were those developed during+ the _ational SST Proqram. Other criteria

were considered but found to be less _atisfactory, For hiqh speed cruise and

landing _pproach (normal operation), modifications to the SST Time Response

o

+I

CriLeria, which were based on the Shomber-Gertsen Criteria, . re found to

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adequately define the handling qualities results of this study. Results {rot

the landing approach (minimum-safe) study were found to be best defined b, the

Pitch Divergence C:iterion established during the National SST Prograr:, Tree

stall recovery control power study has resulted in a new criterion in terns o f

nose-down angular acceleration capability. This criterion has not been pre--

viously established by quantitative test results.

Continuation of this study is recommended in those areas not covered by

this study and in those areas where unanswered questions exist. These areas

are as follows:

o Hiqh speed cruise maneuvering with a simulator having substantial

greater load factor reproduction capability

o Stall recovery with varying stability levels at stall

o Landing flare

o Effect o{ structural modes

o Climb, cruise and transonic speed stability

Future work in the area of handling qualities criteria developmert should

utilize _ generalized math model that includes nonlinear characteristics,

speed dependent derivatives to represent the effects due Lo changes in r.!ach

number and airspeed, and structural modes. Also, it would be essential to

provide the capability to control all of the above and the aerodynanic

characteristics easily. Such a math model would allow , more e_:_icient nroqrar.

to be conducted.

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I''" 2

2.0 I tITRODUCTIOH

This document presents the results of the piloted si,iul(_tion study _,onduLted

under NASA contract (NAS_P-7966) "Development of Longitudinal Handlin_j _,_ualities

Criteria for Large Advanced Supersonic Aircraft." The purpose of this study was

to improve the data base and handling qualities criteria for large supersoPic

cruise aircraft with highly augmented flight control systems,

Research work conducted during the National SST Program has showp that

important benefits in aircraft economics will be gained through advancements in

flight control system design. These advanced flight control systems character-

istical]y result in airpla-e t'-.'namic response which is not adequately specified

by existing handling qualities crlteria. Existing military (Reference l) and

civil (Reference 2) flying qualities criteria were found inadequate to provide

design gu;dance for the flight control system development o_ the large, low

design load factor, SST aircraft. An extensive set of criteria was developed

and documented (Reference 3) during the National SST Program which was based or:

previous work done by NASA and other investigators as well as extensive contrac-

tor fixed base in-house simulation.

Generalized criteria are required for flight control system design a_idance

for both normal operation and minimum-safe operation. For normal operation,

these criteria will establish control syst.em design requirements, augmentation

system requirements, and the requirements for control surface rates and

authority. Criteria for minimum-safe operation are renuired to establish

minimum stability levels and key elemencs in the basic air;)lane design such as

fore and aft limits of longitudinal balance and tail sizinq derived From control

power requirements.

The piloted simulation study covered in this report was conducted usimq

the NASA-Ames moving base simulator designated as the Fliqht Simulator for

_a

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Advanced Aircraft (FSAA). lwe simulation lesl _eriods were utilized, ,_overir,_#

the periods from May 15 throuqh June 6, 1974, and from September II [I,rou,]h

October 25, 1974. During these two periods there were 61.7 hours of pilote_

evaluation time utilized. In addition to piloted evaluation, these periods

were also used to conduct the necessary checkout work and do the required test

set up and calibration work.

This study contract covered the time period from January 15, 1974 through

March 31, 1975.

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3.0 STUOY AREAS There were four basic study areas investiqated during this piloted simulation study: I. High Speed Cruise Maneuvering 2. Landing Approach Inorma] operati(;n, _ 3. Landing Approach Iminimum-s_fe operation) 4. Stall Recovery Control Power These represent the most important Droblem areas in terms of lon_;itudina_ handling qualities criteria identi{ied during the _dt_cna_ SST Proqranl. Also, the selection of the study areas was in¢Ijer, ced by the _:).<]icabilit,,. to each study area of the type of evaluations i)ossiblc_, the faLilitv beinQ used, and the availability of evaluation test time. Fo,- e_<_, F,i, =. these studies did not include, evaluation where large sustained load factor was required due to the _imited vertical strnke built into the motion syster _ o' the FS#,A simulator.

The results of evaluations conducted in eac. _- stud_ are<_ will be ciscussed under separate heading in this report. Crite_1_ a',r4ved ,,_t in each study area will be identified in the discussions o': t_le test resu',ts and _,ill be summarized in the conclusions.

I RE V SYM .. _B__°_ ._N_°__ 17 _i, ......

a.O TEST FACILITY

The facility used for all evaluation testinq was the Flight Simulator for

Advanced Aircraft (FSAA) at ,_IASA-A_Ies Research Center. This simulator consists

cf a large cab with two crew stations mounted on a six decree nl _reedom motion

system. A visual system is provided at each ;Bilot station by means r_f a color

television system using a terrain model disI_]_v. The. ¢,ntire fa_ilit., is

controlled by _ Sigma 8 computer which, in the case _)f thi'. _,i,'_ulation nvaluation_

d

was orogrammed with the complete math model representutior ,;' t'_e ["._.,,._ir_,; 2707-300PT Supersonic Transport (Reference ._.].

Two cockpit configurations were used for thi_, ,_v,_l_ation. These confiqura-

tions differed in the attitude display instrument. As shown in Fi!)ur,, &-l, thp

mechanical attitude display indlcat,.Br (HZ-6F) was ,_sed _urin(. the first simul_:tlerl

study period and had a pitch attitude display ser.sitivitv c,_ .07 InLhes per

degree (.18 cef,timeters per deflree). The other conti,)uration presented in

Figure 4-2 utilized the electronic attituded|rector indicator ([ADI) which was

developed for the 2707-3OOPT durinq the ;Jational SST i'r_,_rav_. _k_s was "he

configuration for the second simulation study _)eric, d, a,,d hat! a normal [itch

attitude sensitivity of .16 inches per degree (.41 cent r_eters p_r deqree_.

For hiqh speed cruise evaluJtions the sensitivity was increased to .B0 inLk L_er

degree (.76 certimeters per degree_j. The mechan'::al ".[" was Jsed to aCcomplisM

all of the landing approach (normal operation) evaluati )r_except those cases

evaluated by Pilot "F". Th_ EADI was utilized for ,_',] _c the other evaluations

and the details of this display are presented in Fiqure 4-!.

A flight path anqle display was available to the I.-Ir, t in either cockpit

configuration, The display consisted of actual an,'. _'.t_ntia_ _li(;_,t _;ath angle

indicators. With the ADI configuration the angles were presented on two

' 6

t

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O _ Q) Lt- Z r_ I ILl

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LL_ I---- t--

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I-- ,::I2 L,J __D L1: !/ _D CD c_ _D _J klA __1

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r' ._Lsl LL .--1 .-J , 4 I " I :|,I ' "2' 1: D

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adjacent vertical scale Instruments located to the immediate right of the ADI

(Figure 4-1). With the EADI configuration these angles were displayed directly

on the EADI display sclope (Figure 4-3).

The purpose of this flight path angle presentation was to aid the pilot

in stabilizing the aircraft. When the potential and actual flight path angle

are the same, the aircraft is neither accelerating nor decelerating. When the

potential is less than the actual flight path angle the aircraft is decelerating

and vise versa. With this additional presentation, thrust management, particu-

larly at high speed, is much improved and does not detract from the primary

task of longitudinal handling.

For a more detailed description of the technical aspects of the test

facility, refer to the section of the Appendix titled "Simulation Facility

Description"

This _acility did prove to be a very useful tool in conducting this tyoe

of evaluation. Numerous pilot comments were received that favored the movinq

base feature over a fixed base simulator due to the added realism. The added

realism was especially noticeable when attempting to stabilize the aircraft where

the small variations in load factor were an aid to the pilot.

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5.0 TEST AND ANALYSIS _ECHrlIE)UE

i i

The objective of this study was to develop handliml qualiti___ _,-ite,-i _ '_',-

large advanced supersonic aircraft. The criteria were to be develo_ed in ter_:s

of airplane response characteristics with primary emphasis on the lonqitudinal

modes of motion. All experimental results were based on the results of piloted

simulation evaluation using both Boeing and NASA test pilots.

The basic approach taken was to identify study areas where criteria

development was important to future desiqn concepts_ and compatible with a

piloted evaluation study usinq the FSAA. For each study area critical airplane

response parameters were identified based on previous experience ,:ained durincl

the National SST Program and follow-on SST studies. Each parameter was then

varied in a systematic manner holdin(_ all other parameters corstart or near

constant, and piloted evaluations conducted. For some study areas the e_fect

of control force gradient, atmospheric turbulence, and Ditch attitude dis;-_l_.

sensitivity were also evaluated as contributinq parameters tc the handlinr:

qualities criterid. The iJilot evaluations nf each test rendition were done

while flyin_i the simulator through a spn_ifi¢ sequence of tasks which were

standardized For each study area. The p_lots then rated each test (:onditir,n

using the Cooper-Harper ratin,_ scale (Reference 5) for th_ h_indl!n,; ,:,jalities

rating, ar, d a turbulence rating scale (Reference 6) for the cases _n_:!vir.:

turbulence. The pilot also provided comments to sFecific cuestions which _,ere

standardized for each study area. In addition to these evaluation ratinqs and

comments pro,'ided by the pilot, data were obtained consi_tinq of pilot deszrib'r.,; function measurements, r_ilc_t workload measurements, and pilot perforr:anc,:, measurements in conductinq the tasks.

The parameters that were varied For the nurnoS_ of this evaluatior ,.._:l be

r

described in the discussions of the test results coverin:_ eacm studl a.e_. '-'atm

• L

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model configurations required to achieve each specific response charactet_ist_c

were defined during engineering calibration runs prior to the piloted pvaiuations i

These calibYation runs were achieved by measuring the lonqitudinal airplane

i

response to either a column step or pulse command while making changes to t_e

math mo_el in the followin(l areas:

o longitudinal SAS t.lain

o longitudinal SAS time constants

o forward loop column prefilter

o airplane center of gravity

o airplane lonQitudinal moment of inertia

o additional tail lift and pitchind moment increments

All evaluations were made usinq the math model of the Boein_I SS- confi:uF-

alien, the 2707-300P_- The math model representation of this aircraft is

described in Reference a.

5.] PILOT RATIt, IG SCALES

Pilot ratinqs were obtained using the Cooper-Harper ratinu scale (Re_erence

5) for the basic airplane handling qualities, and a turbulence rating scale

(Reference 6) to describe the effect of atmospheric turbulence.

The Cooper-Harper ratin_l scale (Fiqure 5-I) was used by all pilots to

describe the lonqitudinal handlinq Qualities immediately after conductin,) Lhe

specific pilot tasks which were standardized for each study area. Two lines of

division were established on this ratinq scale to define normal ooeration and

,linir;ur_-sa+e ol_eration lir_its, lhese limits are the same as used durinq tne

;_d*ional SST Proqram and have been universally accepted. The limitinq pilut

rating 'mR) for normal operation tests was estaL.lished as a ratinn of 3.5.

This ratinq is the dividinq line between a ratinq that requires no im_rove_.ent

(PR=3.0 and one that does warrant improver,ent (PP,--4.0_. Therelrore, ,._aracter-

.....................

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DEMANDS ON TIlE PILOT PILOT ADEQUACY FOR SE_..ECTI;:D TASK OR AIRC;RAI_ CHARACTERISTICS IN SELECTED TASK OR REQUIRED OPERATION* RATING REQUIRI_D OPERATION _

D

_0_ dec,s_ons F_gurc, 5 I. Handling Ouaht_es Rating S_.ale r_ REV SYM

L

istics thaL result in a pilot raLing of 3. C or better are said to have s_tis-

factory handling qualities for normal operation.

The other dividing line was established in a like manner at 6.S _or

minimum-sa{e operation. This is half way between a r_tin[i that describes barely

adequate performance and tolerable workload (PR=6.O) and one that describes

inadequate performance and an untolerable workload (PR=7.0). Therefore, a

characteristic that is rated 6.5 or better is considered acceptable _or m_niT',u_"-

safe operation.

The turbulence rating schedule (Fiqure 5-2) describes th_ effects atmos-

pheric turbulence has on hand]inq qualities and pilot workload. As in tne case

with the Cooper-IIarper scale, a boundary has been selected that represents the

dividing line between acceptable versus unacc2ptable ratings. This re_resents

the dividing line between a rating that describes a confi_uration where all

tasks can be performed and one where som_- tasks cannot be performed. 7 rat!no

o _ "F" or better represents a condition that is acceptable. A ratinq c{ r'G

or worse represents a condition that is unacceptable.

5.2 'ILOT MATH MODEL

[he purpose for determininq the pilot math model was to sunport the

understandinq and interpr_,tation o{ the pilot ratinq data as well as advance

the state-of-the-ar_ in this area. The apnroach taken was to develoo a method

whereby the pilot rating trends could be predicted based on the nilct describinq

function alonu with additional performanc_ and workload Parameters readily

available _rom the experimental data. In this way configurations resultinq in

larqe oi_,ot ratin,l scatter could be re-evaluated based on the pilot describinc

function technique and an indication of the best data {airinq obtained. This

approach was successful to a very limited degree in the landing dPDroach

(normal operation) study area and was the only study area where this approach

was used.

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II4CREASE OFPILOT

I)ETERIORAT]OFi OF [AS;, RAT I NG

PERFORMANCE _#ITH

EFFORT I_ITH

TURBULENCE TURBULENCE

.... I "'_"

NO SIGNIFICANT NO SIGFIIFICAHT

INCREASE

DETERIORATIOFI

NO SIG!_IFICAHT

DETERIORATION B

MORE EFFORT MINOR C

MODERATE D

REQUIRED

MODERATE

NAJOR (BUT EVALUATIOII

BEST EFFORTS TASLS CArl STILL BE

REQUIRED ACCOMPLISHED)

LARGE (SOME TASKS

CANNOT BE PERFORMED

UNABLE TO PERFORM TASKS

FIGURE 5-2 - TURBULENCE[ EFFECT PAT[FIG SCALE

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Data analysis by this technique requires extensive auditional engineerir_;

effort in both data recording and set up requirements, as well as data reduc:tion

and analysis effort. Also, special oiloted evaluation runs wit_ _;di_{er,_nt

but similar pilot task, wererequired to obtain the necessarydata _,o_analysis.

This technique was applied only in the landing a#proach(normal operation'j

study area. This wasdictated by a reduction in available evaiuatinr test tir_e

over that originally planned and the nee_ to conduct higher _riority Lestinq.

A description of this analysis technique is included in the a;_pendi× of

this document, and includes the theory bei_indthe approach, the calculation

techniques, data handling, and analysis techniques. The correlation between

the results of this analysis technique and the pilot rating data will be

discussed in the appropriate section of the discussion coverin_ test results o_

the landing approach (normal operation) study area.

5.3 WORKLOAD MEASUREMENTS

The measurement of pilot workload was accomplished by two different

techqiques. One was by the use of a side task for which the performance could

easily be measured, and the second was by inteqratinq the work nroduced bv the

pilot through column deflection over the test ru_,.

Measuring r_iiot workload b v the use of a side task is accomplished by

measurin_l the performance of the pilot in performing the side task. In theory,

an increase in _ilot workload ir Derforr;_in_ the primar.v task, _lyi_- the

airplane_ will result in a decrease in his perfor,::._nce in accomplishinq Liq_, side

task. This should be true if the side task is considered by the I,ilot to b_

only a side task that is to be accomplished on a totally non-interference Lasis

with respect to the primary task.

The side task selected for this particular study was a li,lht cancellir.':

task. Three li,lhts were located in the cockDit, oroqrammed to come en in a

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random fashion. The pilot was tu turn the lights off when they did corse on by

pressing the 11ght flxture itself. The performance measurement of this task

was made by averaging the time the lights stayed lit. The longer the ti:r_

duration, the poorer the performance of the side task.

This approach to a side task was believed representative of a normal side

t

task that would occur in an aircreft cockpit which the pilot would need to

perform during an actual flying situation. Location of the lights was selected

to support this idea. One light was located directly in front o _ the _ilot on

the glare shield. Another was located at the far left of the instr-ur_:t _a_iel,

on the window sill, and the third was located on she aisle stand immediatel_

aft of the throttle quadrant which was just at the edge of his _eripherial

vision. These locations covered the full range of visual scan normally

maintained by the pilot during the piloted tasks beinq flown.

This approach to measuring the pilot workload was not successful. Jests

P

wheYe the workload was obviously increasing, such as in the case with increased

turbulence, resulted in a decrease, in some cases, in the averaqe t_me the

lights remained lit (Figure 5-3). This was exactly opposite to the exnected

P

results. The main reason attributed to the failure of this approach _as the

color of the lights used. The lights were amber, which usually denotes a

significant malfunction in the cockpit. The attention, therefore, qiven to tn _

side task was higher than desired. With an increase in work load, the pilot

worked harder to keep the lights off in order to minimize his distractions.

Plans were implemented to repeat this evaluation of the workload measure-

t

ment technique using the side task with different colored !i_hts, such as blue.

Chances of success with blue lights were believed to be much _Ireater s;_ce it is

a color that the pilot is not trained to look out for and react to. _owever,

this was not done due to a shortage of simulation evaluatiun ti:_e ard hiqher

priority work that needed to be accomplished.

i

BOEJN_ __

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I

@

• -w_ x _ \._ .S_c_O___ G__/kC:t qmC>V,V t'-4 _'k../k_S = ?_-0 _ ,,01 .0 -

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o 8

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APPO _I G. 5--_

W _,A"_,-X -_ t___:_:A._ X__ _.___'-i C ____ APPO • ._ _9'_TEl/t/a_' co,,..,,.

t IH_,<NI AO 4ill ¢

The second method for measuring pilot workload was to integrate the work

the pilot does through the control column over the test run. This is representec

by the following formula:

COLUMN WORK LOAD = (Fco I x _cel) dt

This approach did represent the physical work the pilot was required to

perform and was used to judge the validity of some of the pilot ratinqs,

particularly in areas of large data scatter. This will be pointed out in the

specific portion of the Study Results section where the workload data v_as

successfully used to help interpret the pilot rating data.

5.4 WIND MODEL

The wind model was based on information contained in Reference 7. f,

summary of the wind model parameters is also contained in the appendix of this

report.

Evaluations with the wind model were conducted for all but the hiqh soeed

cruise evaluations that included the landing approach for normal and minimum-

safe operation and the stall recovery control power tests. In all cases the

-3

maximum turbulence level was based on a probability of exceedence of lO which

defines a crosswind velocity oF 25 knots (12.86 meters/second) at the reference

height of 20 feet (6.1 meters). For the landing approach (normal oneration)

study and the stall recovery study, a terrain roughness factor (_g) of l.O was

used which defines the roughest terrain expected in the vicinity nf any a_rport.

With the combination of the wind velocity and roughness factor, the wind model

produces a root mean square vertical turbulence comDonent !C w ',. _dlue 0 _

approximately 7.0 fps (2,13 meters per second). For the laK1ding aD_)roac _

(minimum-safe operation) a terrain rouqhness factor cf .15 _as used which

Justification for using the ;ower

represents the normal airport terrain.

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roughnessfactor was based on the reduced probability of the mini.lu.i-safe

configuration occurring. All probability estimates used in establishin_l th_

wind model configurations for each study area were obtained from probability

studies conductedduring the National SSTProgram.

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.__ -

_J I

i

6.0 EXISTIrlG CRITERIA COMPARISON

I

The purpose of this study was to develop a handling qualities data base of 1

I

aircraft response characteristics that will improve the data base of design 1

criteria for large advanced supersonic aircraft. Emphasis was placed on 'qe

longitudinal axes since this is the area of greatest benefit in terms of

increased airplane efficiency.

Comparisons against existing criteria were done where oossible in order to

substantiate, modify or de-emphasize the criteria established by previous

studies. Where this approach was not appropriate new criteria were established.

By accomplishing the analysis in this manner the strongest data base for long-

itudinal handling qualities criteria were believed to result for those areas

investigated during this study.

The greatest source of existing longitudinal handling Qualities criteria

for this type of vehicle has been from the National SST Program. During the

National SST Program the sensitivity of airplane design to criteria variation,

and hence the need for adequate criteria, was demonstrated. A considerable data

base was established during that program since existing criterla at that time

were found to be inadequate for ma×imum design efficiency of a vehicle of this

type. The criteria from this data base used for comparison purposes were the

SST Time Response Criteria for normal operation and the Pitch Divergence

Criterion for minimum-safe operation. ()ther (riteria used for comparls.),;

purposes consisted of the C* Lor:_itudinal llandling Qualities Criterion and

applicable criteria rron the military handling qualities specigication, MIL-F-8785B.

The SST Time Response Criteria were obtained from Reference 3. These

criteria are in the form of time history envelopes in response to a ste_ colur:_r

input. Pitch rate and nori,al load factor are the terms used to define these

I

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I

time history envelopesas seen in Fiqures 6-I through 6-3. Theseenvelo_Jes and

these criteria are basedon the Shomber-Gertsen Criteria. The Shomber-Gertsen

Criteria arebasedon the results of several other studies and is expressedin

terms of L_/_J-_ and_/_versus dampingratio (_). By exl_ressingthe resa'ts

of the other studies in these terms the results were found to converaeinto

common boundariesas detailed in Reference8 and presented in Figure 6-_.

The problemwith using the Shomber-Gertsen Criteria directly is that they

arebasedon a simple secondorder systemand direct comparison with higher order

systems would be inappropriate. However, such a comparison is possible by

comparinq the time history response to a common input command such as a colump

step of the second order system to the higher order system. This was the

approach taken durina the _lational SST Program in develo_)ing the ]onqitudinal

response time history criteria which will be referred to as the SST Time Response

Criteria in this report.

The pl lot rating scale used in the Shomber-Gertsen Criteria was the Cooper

ratinq scale (Reference 8) [)resented in Figure 6-5. This ratin_l scale is a

simplified version of the present Cooper-llarper rating scale (Reference 5),

presented in Fi,lure 5-1. Results obtained with this earlier scale are believed

comparable with results obtained usinq the later scale. The critical dividiro

lines are the same. Thal. is. between the ratinqs 3 _nd _ the airplane handlinq

qualities chan,le fro_ sa_.isfactory (indi(.atinc! no it,lprover',ent necessary) to

unsatisfactory, and between the tatings _ and 7 the airplone handling qualities

change _rom acceptable Lo unacceptable. These are the same critical judgments

made during this simulation study.

The C_ criterion was deri4ed by Boeinrl using fliqht test results from the

Cornell Aeronautical Laboratories (now known as <alspan) variable stability

] 367-:;0 variable stabil4ty airplane (Reference 9).

., air[_lane and the Goeinq t_ode

,IThe criterion consists at a term desi,lnated C* which is comnuted usi_q both

L__

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_.>_

,Quarry

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L

COOPER RATING SCALE

ADJECTIVE DESCRIPTION NUMERICAL

CATEGORY WITHIN CATEGORY RATING

L',,

ACCEPTABLE

EXCELLENT

AND GOOD

SATISFACTORY FAIR

ACCEPTABLE FAIR

BUT POOR

UNSATISFACTORY BAD

BAD

rl

UNACCEPTABLE VERY BAD

Q

DANGEROUS

UHFLYABL[ 1O,

ADDITIONAL DEFI_ITIOqS OF UNACCEPTABLE CATEGORY"

7 BAD - AIRCRAFT COtlTROLLABL[, BUT REQUIRES HAJOR RORTI@_I OF PILOT'S

ATTENT [O_i

R VERY BAD - AIRCRAFT CO_TROLLABL[, BUT ONLY WITH A MJ_IMUM OF COCKPIT

DUTI ES

9 DANGEROUS - AIRCRAFT JUST COIITROLLABLE WITH COMPLETE ATTE!ITIO!;

10 UNFLYABLE

FIGURE 6-5

JD'dI_EJ_Vd_:' ! ".

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L

• rl 1111111 _ I --q

pitch rate and normal acceleration. The time history of this term in response

to a column step input must fi' inside a specified time historl resnonse

envelope For the handling qualities to be acceptable (Fiqure 6-6].

Extensive work was done in the area of minimum-safe handling qualities

criteria. The purpose of such criteria is to define the allowable deqradation

of handlin(l qualities (as could result from stability auqmentation syster-

failures) that would still provide safe handling qualitips under erlerflercy

situations. The primary area of importance for this probler is landinc apprc.a_n,:

which was one of the study areas selected for this study cortra_.t. The mir. imu_'-

safe criteria established during the ,_lational SST Proqram were based on the rate

of pitch divergence as the result of a momentary pitch disturbance such as a

longitudinal c;ulse Inuut. This same aFJDroach was taken durin_I this study, and

a typical response is presented in Fiqure 6-7. The diver(Jence rate detert;:ined

acceptable durin!l the National SST Program was a time-to-double pitch a::_plitude

of 6.0 seconds. The diverclence rate criteFion from the ,_ational SST ProQram was

based on the most unstable root. Basing the criterion on the most u_stdble reel

was done _or ease. of control system desiqn. During this simulation stud)' the

r,eti:od for ,:easurin,: the divergence rate is believed to result in a reasuremert

of the T'.:ost unstable root. This allows direct comparison with the '_ational

SST Pitch Divergence Criterion.

The military specilcication, tIIL-F-;_,7_$SB, was examined in terms o _ the

sl)ecified li,,its of short l_eriod natural frequency, l)itch d_mpinq ratio and

column force gradient, llowever, the criteria in this specification have beet

based on i,aru,_eter variations of a second order system which are not directl_

applicable to the results of this study. This study utilizes an actual nor.-

linear a_ro'_ne 1,_ath model and au_Imentatior, sysLem which results in :_itch

res;mnse characte_istics o _ a :'_uch hiqher order than a second order s;'steF.

This '_tudy, as well as the _iational SST criteria development study, was ai_ec

- .....

i i i L _'_..C_'_ "_l_."_'or_-,,. ":5._ I m I CALC CHECK C. "_ V_Q,_r_ _._ _._;_\'_.__,\_'rA APPD APPD r , ,

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at developinq a criteria data base that could be applied to aircraft that

had predominantly higher order pitch response characteristics.

I

P

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7.0 S[/U DYRESULTS

This piloted simulation study was broken downinto separate study areas each

covering a particu]ar flight regime or characteristic, The results of the work

done in each study area will be presented separately along with a description of

the test techniques and a description of the analysis used.

7.1 HIGHSPEED CRUISE MA_IEUVERING

tti{;h speedcruise presents a different set of handling cualities require-

ments than does low speed, This is mainly attributable to the larqe true

velocity vector and the resulting sensitivity of rate of climb or descent to

lonqitudinal attitude and rate of changeof attitude, Thesernquirementsare

reflected in the pitch attitude display sensitivity requirements, the desired

longitudinal responsecharacteristics, and the importanceof load factor as a

short period parameter, i

Tl_efliqht condition for this evaluation was the condition occurrin!l at the

end of supersonic climb for the 2707-300PT airplane, identified as follows

o l,lach2.7

o 60,000 feet altitude (18,288 meters)

o 567 knots CAS (292 meters per second)

o 555,000poundsqross weiqht (251,744 kilograms)

o 62 center of gravity (aft limit)

The parametersidentified for evaluation of this study area were the short

period responseparametersmeasuredin terms of pitch rate and expressedin

the following terms

,, o pitch rate overshoot ratio, @max/@ss

o time-to-peak pitch rate, T@

7 max

- o pitch damping! constant, _

I'" " 32

$

These expressions define the pitch axis short period response characteristics

_valuated and are the same parameters evaluated in the landing approach

(normal operation) study area. These same short period responses were also

analyzed in terms of the normal load factor response characteristics.

The response characteristics just identified were developed accordi_ tc

the technique described in the "Test and Analysis Technique" section of this

report. The forcing function for the engineering calibration runs was an

unpiloted column step input.

In addition to the short period response parameters, the effect of varia-

tions of the column force qradient and pitch attitude display sensitivity were

also evaluated. Table 7-[ presents the complete matrix of the parameters J_ed in

the evaluation of this study area. Time histories for the short period resmense

parameters in terms of normalized pitch rate are presented in Figures 7-1 thrcuqh

7-3. These same short period response conditions in terms of normalized

load factor response, are presented in Figures 7-4 through 7-6. Both sets of

time histories have been normalized to a steady state value of unity.

Results of the evaluation of each parameter will be discussed separately ir

the following sections. Under each section the results will be compare_ dc]ainst

the existing criteria listed below:

o SST Time Response Criteria

o C* Longitudinal llandling Qualities Criteria

o Military Specifications (MIL-F-8785B)

The pilot task description for evaluation of this study ar_a is t_resented

in Figure 7-7.

Z

., 7.1.1 Pitch Attitude Display Sensitivity

%

For all testin_ in this study area the electronic attitude director

; indicator (EAD[) was used (Figure 4-3). Variation of the pitch attitude scale

I

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I,-' ,, 33

im ii

TABLE 7- I

HIGH SPEED CRUISE HArEUVERING TEST CONDITIONS

PARAMETERS NUMBER OF PILOTS

VARIED EVALUATING (Smooth Air)

EADI PITCH SCALE

in/deg = .]6

.23

* .30

PITCH RATE OVERSHOOT RATIO

Omax/@ss = 1.94

* 4.10

6.10

,q.2

TIME TO PEAK PITCH RATE

TA = .45 sec

t_

max * . 80

l .40

2.0

DAMPING CONSTA_IT

.36

.90

2.42

COLUMrl FORCE GRADIErlT

lO ]b/q

Fcol/q

* 25

_5

Baseline confifluration

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• lOGO AD 4till ¢

L

r

HTGH SPEED CRUISE MANEUVERING

PILOT TASK

ALTITUDE CHANGES (HOLDING MACH NO. CONSTANT)"

i. CLIMB 250 FT @ 500 FPM (76M 0 152M/MINUTE) AHD STABILIZE

2. DESCEND 750 FT @ 1000 FPM (229M O 305M/MIHUTE) A_D STABILIZE

3.

CLIMB I000 FT O 2000 FPM (305M @ 610M/MINUTE) AHD STABILIZE

,_.

DESCEND 500 FT @ 500 FPM (152M @ 152 M/MIHUTE) AriD STABILIZE

AIRSPEED CHANGES (HOLDING ALTITUDE CONSTANT)'

I. INCREASE SPEED 20 KNOTS AND STABILIZE

2. DECREASE SPEED 40 KNOTS AT',IDSTABILIZE

3.

INCREASE SPEED 20 KNOTS AND STABILIZE

HEADING CHANGES (HOLDING ALTITUDE AND AIRSPEED CONSTAHT)

I. TURN 150 LEFT IN 15 o BANK AND LEVEl_ OFF

2. TURN 200 RIGHT IrJ 30 o BANK AND LEVEL OFF

FIGURE 7-7

BOIfNG I _,.,

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was desired as part of the pvaludti_n ,,_ this studv are_ since d rlredter sers!-

tivity is required at hiqh supersonic cruise speedsthan at subsoniL _r,ise

speeds. The l_itch attitude scale sensitivity requirement should be rouqhl/

proportional to the magnitudeof the true velocity vector which defines the

relationship betweena changein vertical velocity and a changein pitch

attitude. For example, one degree of pitch attitude at !4ach2._ r'esults i_ 2___.

feet per minute vertical velocity. At Mach.8 one degree of pitch chamc_e ,es_]ts

in approximately 800 feet per minute _ertical velocity. With the recuire,'e,'t

established for a qreater I_itch attitude sensitivity, the objective was to first

define the optiHum pitch attitude sensitivity and then conduct all other evalua-

tions at that scale sensitivity value.

Three pitch scale values were evaluated as seen in Table 7-I. The results

ol this study _re presented in Fi. ,_e 7-_). Both pilots preferred the ,30

inches/de,l. (.762 centimeters/deq) sensitivity accordinq to the ratings qiven

and according to their comments. They were both given their choice of any c_

the three settings for the remainder of the e_aluation o_ this study area, and

both selected the most sensitive setting. Also, it should be t_ointed out that

this ,_as the f;ost sensitive setting possible with the EADI svste" avai]ablr_.

This wa: due to the spacincl c _ the pitch bars approachinq the lirit of the

screen size available. Pilot comments were received durinq evaluation c ,_ othor

parameters indicatinq a _ore sensitive r,,itch scale would be desirable.

7.1.2 Pitch Rate _lvershoot Patio

]he data are i_lotted and faired in l i._ure 7-_. The rairin,; selected Srlo_'vs

the o_tir_u_, _a]ue of over,hoot ratio tu be less than _qe lowest value tested o'

i

".i i.9C. Therefore, a lower lir",it could not be established _ro,' these tests. -_e

upner li:it, tmwever, _s shown to be _.i, w_.ich cor:_I)ares fa._:r_r_ly _ith the

',_e,lk value ,,* 7 (, obtained rror the c,-- i;i,;h Speed Pitch m,_÷_ _r,S,Orse Cr _*_r'_

I

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r To_,.,,uO " : ,O.9 .S_C..

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(FIHure /-10). llow_'vnr, (i,, s_,(.'n ir_ thi _, fi,lure, the cmit(.rion shows a ti,,_, to

reach thls overshoot ratio as .5 seconds, and for these tests the time to teach

an overshoot ratio of 7.1 can be interpolated to be .9 seconds, which could

account for the slight difference in the satisfactory ma×imum level. By

comparing this value of overshoot ratio and time-to-peak value a!_ainst the tlr:_e

history envelope of the SST High Speed Pitch Rate Response Criterion, it carL be

seen that the limiting response condition would tall just slightly outside the

boundary. Based on the scatter in the pilot rating data, and the fact the

described differences are very slight, no change is recommended in the ui_oe_

limit of the SST lliqh Speed Pitch Rate Response Criterion ti_:e history envelope

as a result of this test.

On the low side of the envelope, however, the SST Hi_!h Speed Pitch _ate

Response Criterion is not satisfactory. Considering a time-to-first peak o _

.9 seconds, as was used for the overshoot ratio tests, the ,:inimu_:; acceptable

overshoot ratio should be lowered. How low it can go and still be satisfactory

is unknown at this time. However, since steady state pitch rate values for a

given load factor are quite low at this '.lath number, it is logical to assume

that pitch rate overshoot ratio becomes unimportant at ]ow va]ues. Therefore,

no .iustificati_n is api;aren_, fo_ _'_quirirLq _J minimur_ value greater t_,_r_ 1._

(i.e. no overshoot). Therefor(:, a recom_,_ended boundary r,odificatic, r. is tc

truncate the, lower b_bndarv at an overshoot ratio o¢ i.O. The ini/',:" resT;;_rsc,

rate at low overshoot ratios becomes the irrportant parar.leter, and will be

covered in the next section.

?, _.omparison with the SST l;iqh Sipped [.oa(_ Factor Pes[,onse Criterion ¢o,"

the sa,"e set of responses .iust discussed is presented in Fi,:.,,-_ 7-11. T_, loac

factor res_)_rses do not compare favorably _Jl_h th_ criterion. "his :so,"

comparison could be attrihuted to the fact that the. systeT e_a:_ateG is ,.: "_.'.m

represented by the cFiteFior. Thi_

hiqher order than the second order one

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"7 C-_

_q z

I I II _ | | _ | - 1 A_P_ APP')

I

Q _I00_ H Z,c_ d G; .J _J -.% CALC CHECK F\G. "7 - %% APPO APPO PAGE

I

particular set of response conditions does not behave like a second order s_ste,, i

The pitch rate time history envelope comparison is much better for this set e _

conditions.

Another point to be considered when analyzing any load factor response data

is that the motion system used for this study had very limited vertical tra'.el

(+ 4 feet or + ].22 meters), and therefore, limited load factors reproduction

fidelity. This limited capability could account for the poor comparison with

the criterion. Therefore, it is recommended that additional work be conducted

in the future in tile area of high speed handling qualities usinq a mevinq base

simulation with greater load factor reproduction capability.

The military specification (Reference 1) criteria, expressed in terms o _

natural frequency (_) and n_ , is compared against the results of this test in

Figure 7-12. As can be seen, all points fall within the satisfactory range, even

the point for an overshoot ratio of 8.2 that received a pilot ratin_I of a.O.

Aqain, this criteria is based on a second order system whi,:h is not representative

of the system evaluated.

A comparison was also made with the C* criterion and is presented !r,

Figure 7-13. This comparison is for the baseline configuration which was <ive_

a pilot rating of 2.5 and 3.0. Since the C* response con'muted violates the

satisfactory boundary, there is disagreement between this criterion and previous

conclusions. In this portion of the fli_lht regiue the reason is attributed to

the differences between the r,_ of the ,qath r,)odel used in tPis evaluatio_ e,;d

the test aircra*t (variable stability F-94) used in the develol)ment of the

criterion. The 2707-300PT math model has a low n compared to the F-9_

aircraft. This effect is seen in Fiqure 7-13 in the increased rise time c _ L*

over that defined by the criteri n. This increased rise time is primari]# _ue

to the load factor contribution to the C* parameter.

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In conclusion, the modified SST lligh Speed Pitch Rate Response Criterlon

is the best criterion for judging satisfactory pitch rate overshoot ratio

characteristics.

7.1.3 Time-to-Peak Pitch Rate

This parameter gave the greatest variation in pilot rating of all the

parameters tested for this study area. The results are presented in Figure 7-14, and include the estimate of the satisfactory limit.

As was the case with the overshoot ratio parameter, no lower acceptable

limit could be determined from these tests. The lowest limit tested was ._5

seconds, which gave the most satisfactory pilot ratings and compares favorably

with the SST High Speed Pitch Rate Response Criterion (Figure 7-]_). Lower

values are not believed to be of practical interest for larqe aircraft due to

other aircraft parameters such as pitch inertia or structure modal character-

istics causing the predominant restrictions. Therefore, the upper lit:it is the

only concern for this evaluation, and is estimated to be 1.2 seconds.

Close analysis of these results indicates the pilot is not evaluating the

initial response, but is evaluating the duration the overshoot exists. This can

be exemplified by comparing the time history curve for the T_ of 2.0 seconds

ma×

(Figure 7-2) which was rated 6.0 and 7.0 with the time history curve for the

°

@max/@ss of 1.9_ (Fiqure 7-i) which was rated 2.0 and 2.5. Both curves have

nearly the same initial response up to the steady state value. The difference

is in the overshoot magnitude and duration. [n this area the SST High S_eec

Pitch Rate Response Criterion is very accurate in defining satisfactory limits.

These test results compare very well with the SST High Speed Pitch Rate

Response Criterion (Figure 7-15). The time-to-peak of 1.a seconds is _ust

outside of the pitch rate time history boundary and is rated slightly u_satis-

factory. The time-to-peak of 2.0 seconds is considerably outside the boundary

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and the pilot ratings definitely reflect this. The only area of any slight

disagreement exists with the time-to-pe_k of .45 seconds. This does slinhtly

violate the boundary on the low side. However, it should be remembered that

this portion of the boundary is definitely in disagreement with the overshoot

ratio test results, and should be modified as recommended in the discussion of

those test results. With this recommended madification to the boundary, the

response for the time-to-peak of .45 seconds will then not violate the pitch

rate response envelope.

These same responses expressed in terms of normal load factor in comparison

with the SST High Speed Load Factor Response Criterion, are presented in

Figure 7-]6. This comparison is reasonably good in that both time h_stories

for condition rated unsatisfactory are definitely outside of the envelope.

The other two conditions which were rated satisfactory are just within or

slightly outside the boundary. However_ the pitch rate envelope criterion is

judged more accurate in defininq satisfactory conditions for this type of

response variation.

These same test conditions were compared against the MIL-F-_785B natural

frequency criterion (Figure 7-17). The results of this comparison cor_pare

favorably with the pilot ratings received. Meeting the Level I boundary is

considered necessary for normal operation. In this respect the T_max value of

1.4 is just inside the Level I boundary, and was rated a.O and 4.5, which shows

this boundary to be optimistic.

In conclusion, the SST High Speed Pitch Rate Response Criterion with the

lower boundary modification was found adequate for judqinn satisfactory levels

of time-to-peak pitch rate with the corresponding overshoot ratios.

7.1.4 Pitch Damping Constant

This parameter has the same basic characteristics as the other parameters

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tested in that only one limit, the luwer li:'_it in this case, was idertified.

The results are presented in Figure 7-18, and show a r]inimur' s_tis_actory lin'it

of _= .55 fo_ a pitch overshoot ratio of approximately 4, and a t_,_e-to-:ea_

of aDi)roxir_ately .9 seconds. Analysis of this data is believed to be ve_-.,

straight forward. The pilot rated the low damping case as u_acce._tah!_ d',e te

the oscillatory nature of the aircraft response. This is verified by the _,i!ot

comments, and is also supported by the SST High Speed Pitch Rate P, es_onse

Criterion, The comparison is presented in Fiqure 7-19. The response for the

cJ_= 36 condition was found unacceptable and violates thr_ envelope on b th

the low and high side due to the oscillatory tendency ol this response. Lll

other measurements ol this particular- response, such as the _Jversnoot _atic_ _nd

time-to-first peak, would h,:_ve otherwise been satisfactory. The other two

responses are well within the envelope and are rated satisfactory.

A comparison o_ these same conditions expressed in ter:::s o_ ]gag _actor

response with the SST High Speed Load Factor Response Critecion is orese_Led

in Figure 7-20. This comparison is reasonably good in that the tir.e hiStOry

curve for the _J'_ = ,36 condition definitely violates the envelope a_d receivec

unsatisfactory pilot ratin_}s. The uther two time histories are ,,los_, to the

boundary with one sliqhtl/ outsido. This envelo[_e __omr,a,-isor. is Sensitive' and

judged not as qoo(] for corrparison as the i,itch r,:_te _nv_lo,)(:.

Comparinq these test conditions with the !,!IL-F_g7£SB frequent, :riter:a

(Figure 7-2]) qives the e×pected unsatisfactory results with al _ data orimts i, the satisfactory region. However, this speci{ication covers dar_pinc se[_aratel',

b_ dei_ininq a satisfactory i-ini_um volue o c dampir.q ratio ! _ ". For this

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particular aircraft confiquration the lowest satisfdctory valur, e _ _ , se_c

by M[L-F-8785B, is 0.30. The minir_:u_': satisfactory value deter":rec _ro- t_e

curve in Fiqure 7-20 was _ 0._4, which Shows _IL-F-_7'_SB t:, be s:i,:htl_

optimistic. This value of ,_ - 0.3a is obtained bv ravin,: t _ -'_"'_" set's-

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factory value of _&_= 0.55 and dividing by _f_= 1.6 tad/see, _h,ch is "Ke .::'_e

of L_r_ at the test conditions on either side of the intercept noint.

Therefore, the High Speed Pitch Rate Response Criter,on v,ith the lone.

boundary _;odification is judged as the best method for predictinc satisfactory

characteristics for the pitch damp-;nq response parameter.

7.1.5 Column Force Gradient

The results of the colur, n force qradient evaluation arp i_resented it.

Figure 7-22. With the sul_ll amount of column deflection north:all,/ used at the,

high speed cruise tliqht condition ( + 0.3 inches or + ?.q centir_'ters', the

level of the column force gradient is found to be relatively unir'portant. _he

MIL-F-8785B limits are also presenttJ in the above figure for refererce.

Nc specific gradient li_rits are recom:_ended as a result o_ this e_a_ .... .:_

The nominal column force gradient is estimated to be approximately _0 ILs,/q

7.2 LANDINr" APPROACH (NORMAL NPERATION)

The purpose of this secti(m of the study was to develop lonqitudinal

handli:,q qualities criteria based on airplane response char_teristics for the

landing approach portion ol the mission. For this particular landinq approach

study the criteria was to be applicable to airplanes landing} under r, orTYa_

operating conditions as differentiated from airplanes landinq under r'iniru .... sa!:e

operating conditions, as will be covered in the next section.

As described previously in this report, the criteria is based o_ air; line

response parameters. The parameters used to evaluate this study area were

as follows: r .

o hitch rate overshoot ratio (gmax/@ss)

o time-to-peak pitch rate (T_ )

o pitch damping constant ( _ _ )

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The column force gradient is applicable since it is a measure of the feel the

pilot has for the longitudinaJ response characteristics, even though this is

not a response parameter as such.

The complete matrix of parameters evaluated is presented in Table 7-11.

Included in the table are the conditions under which the evaluations were

performed, special tests measurements taken and number of pilots evalbatin(_

each case.

The parameter variations were calibrated using the response from ar

unpiloted column step command. To obtain the desired variations of the above

parameters, changes were made in the pitch SAS by adjusting the gain and filter

time constants and by changinq the column forward loop prefilter terms.

The pilot task is presented in Figure 7-23. This pilot task was I_erformed

by all pilots evaluating the landing approach configurations for both the normal

and minimum-safe study areas. When conducting the landing approach _ilot task

the visual scene was fogged over unti] an altitude of 200 feet (61 meters) was

reached, fit that altitude the fog was lifted and the pilot continued the landinc

through flare and touchdown using the visual scene.

The initial flight condition for this evaluation was the norma] landinQ

approach conditions for the 2707-300PT airplane as follows:

o 1800 feet altitude (549 meters)

o 144 knots CflS (74 meters/sec)

o 415,000 pounds qr(iss weight (188,240 kilograms)

o 54 center of qravity (forward limit)

' 20 deqrees flaps

< qear down