NASA-CR-176355 

19860003826 

A Reproduced Copy 

OF 

Reproduced for NASA 

by the 

NASA Scientific and Technical Information Facility 

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(NASA-Ch-176J55) AN INVESTIGATIOn INTO THE 

Vl-.HTICAL AXIS CCUII1CL FOWEn IiECUIHEHE!lTS POR 

LAUDING VTOL TYFE AIRCRAFT CNEOABD 

NONAVIATION SHIES IN VARIOUS SEA STATES 

(Raosds Univ. C€nte~ fo~ Research, Inc.} G3/02 

tHl6-1J254 

Unclil.s 

04835 

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THE UNivERSiTy of KANSAS CENTER fOR RESEARCH, INC. 

2291 Irving HIli .. Drive-Campus W est Lawrence. I<ansas 66045 I 

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Report on Resear~h 

Performed under 

NASA Cooperative Agreement NCC 2-242, 

"V/STOL Handling and Control Power Requirements" 

AN INVESTICATIOi'I Dn'O mz VERTICAL AXIS 

co:rt'ROL P<mER REQlDIREMEm'S POR LANDmG 

V'J'OL TYPE AIRCRAFT' amoARD NOH-AVIATIon SHIPS 

m VARIOUS SEA STATeS 

KU-FRL-623-1 

Prepared by 

Mark E. Stevens 

Principal Investigator: Jan Roskam 

Flight Research Laboratory 

University of Kansas Center for Research, Inc. 

Lawrence, Kansas 66045 

July 1985 

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ABSTRACT .. 

The problem of determining the vertical axis control. 

requirements for landing a VTOL aircraft on a moving ship deck in 

various sea states is examined. Both a fixed-base piloted 

simulaaon and a non-piloted simula·tion were used to determine the 

landing performance as influenced by thrust-to-weight ratio, 

vertical damping, ann engine lags. 

The piloted simulation was ru~ using a fixed-base simulator at 

N.A.S.A. Ames Research ~enter. Simplified versions of an' existing 

AV-8A liarrier model and an e'xisting head-up display format were 

used. The ship model used was that of a 00963 class destroyer. 

Simplified linear models of the pilot, aircraft, ship motion, 

and ship air-wake turbulence were developed for the non-piloted ! 

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simulation. A unique aspect of the non-piloted simulation was the 

development of a model of the piloting strategy used for shipboard 

landing. This model was refined during the piloted simulation until 

it provided a reasonably good representation of observed pilot 

behavior. Further refinement could lead to a model suitable for 

prediction of landing performance of VTOL aircraft on ships and as 

the basis of control logic ~or automatic landing. 

A surprising result o~ this simulation was that, with a good 

station keeping control system and with statistical ship motion 

displayed on the head-up display, pilots could' consistently perform 

safe landings in sea state 6, ·with handling qualities that were 

adequate at thrust-to'-weight ratios greater than 1.03 and even 

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marginally adequate down to thrust-to-weiqht ratios of 1.01. These 

results should hold quite generally provid~d that a thrust-to-weight 

ratio of 1 + A is interpreted as meaning that the pilot always has 

the capability of accelerating the aircraft at 6g upward even in the 

presence of ground effect and hot gas reingestion. 

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TABLE OF CONTENTS 

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ABSTRACT ••••••••••••••• , •••••••••••••••••••••••••••••••••••••••••• 1 

LIST OF FIGURES •••••••••••••••••••••••••••••••••••••••••••••••••• vi 

LIST OF TABLES ••••••••••••••• · ••••••••••••••••••••••••••••••••••••• x 

LIST OF SYMBOLS •••••••••••••••••••••••••••• & ••••••••••••••••••••• xi 

I. INTRODUCTION .................................................... 1 

I I. ANAL YS IS OF HOVER HEIGHT CONTROL •••••••••••••••..•••••••••••••• 5 

A. COlrrROL TASK ••••••••••••••••••••••••••••••••••••••••••••• 8 

B. THE NON-PILOTED SIMULATION ••••••••••••••••••••••••••••••• 8 

c. A'NAL YTICAL MODELS ••••••••••••••••••••••••••••••••••••••• 1 2 

1. Aircraft Model •••••••••••••••••••••••••••••••••••••• 12 

2. Pilot r~9.del •••••••••••••••••••••••••••••• ••••••••••• 13 

3. Ship Model •••••••••••••••••••••••••••••••••••••••••• 15 

4. TUrbulence Model ••••••••••••••••••••• 4 •••••••••••••• 20 

D. RESULTS FOR THE NON-PILOTED SIMULATION •••••••••••••.••••• 22 

III. SIMULATION EXPERIMENT •••••••••••••••••••••••••••••••••••••• 25 

A. TEST f.1ATRIX ••••••••••• •.••••••••••••••••••••••••••••••• ~ .2S 

B. AIRCRAFT MODEL ••.•••••••••••••••••••••••••••••••••••••••• 25 

1. Basic Aircraft •••••••••••••••••••••••••••••••••••••• 25 

2. Modifications to the Simulation Model ••••••••••••••• 27 

3. Control Augmentation system ••••••••••••••••••••••••• 28 

4. Head-up Display •••••••••••••••••••••••••••••.••••••• 30. 

C. SHIP MODEL •••••••••••••••••••••••••••••••••••••••••••••• 32 -, 

D. SIMULATlorl FACILITIES ••••••••••••••••••••••••••••••••••• 32 

r'-. 

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TABLE OF CONTENTS, continued 

1. Simulator ••••••••••••••••••••••••••••••••••••••••••• 32 

2. Pilots •••••••••••••••••••••••••••••••••••••••••••••• 34 

E. OPERATIONAL "TASK ••••••••••••••••••••••••••••••••••••••.•• 36 

F. RESULTS FOR THE PI~OTED SIMULATION •••••••••••••••••••••• 38 

1. Task Performance and Pilot Ratings· •••••••••••••••••• 38 

2. Pilot Cornments •••••••••••••••••••••••••••••••••••••• SS 

IV. REVISED ANALYTICAL PREDICTIONS •••••••••••••••••••••••••••••• 58 

A. MODIFIED AN~YTICALMODEL ••••••••••••••••••••••••••••••• 58 

B. COMPARISON OF ANALYTICAL AND SIMULATOR RESULTS •••••••••• 63 

V. CONCLUSION'S •••••••••••••••••••••••••••••••••••••••••••••••••• 75 

VI. REFERENCES •••••••••••• , •••••••••••••••••••••••••••••••••••••• 78 

VII •. APPENDIX A •••••••••••••••••••••••••••••••••••.•••••••••••••• Al 

A. APPROXIMATE INVERSE LAPLACE TRANSFORM METHOD., ••••••••••• A2 

B. A.I.L. IMPLEMENTATIO~ ••••••••••••••••••••••••••••••••••• A3 

C. DEVELOPMENT OF THE SHIP MOTIOn MODEL ••••••••••••••••••• A10 

D. TURBULENCE MODEL •••••••••••••••••••••••.•••••••••••••••• A.13 

E. HOVER HEIGHT IUPUT ••••••••••••••••••••••••••••••••••••• A14 

F. DATA PLOTTING •••••••••••••••••••••••••••••••••••••••••• A14 

G. LISTING OF NON-PILOTED SIMULhTION COMPUTER 

VARIABLE DEFINITIONS AND PROGR~M LISTINGS •••••••••••••• A22 

VIII. APPENDIX B •••••••••••••••• ~ ••••••••••••••••••••••••••••••• Bl 

A. TEST VARIABLES •••••••••••••••••••••••••••••••••••••••••• B2 

B. HUD DYNAMICS •••••••••••••••••••••••••••••••••••••••••••• B3 

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TABLE OF CONTENTS, continued 

c. DATA OUTPUT OF PILOTED SIMULATION ......................... B4 

D. PILOTS •••••••••••••••••••••••• e .•••••••••••••••••••••••• • BS ... 

E. LISTING OF PILOTED SIMULATION RUNS AND 

RESULTS ••••••••••••••••••••••••••••••••••••••••••••••••• 89 

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Figure No. 

1. 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

10. 

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12. 

13. 

14. 

15. 

16. 

LIST OF FIGURES 

Non-piloted Simulation Computer Program 

Flow Path ••••••••••••••••••••••••••••••••••••••••••• 9 

Block Diagram of the Original Non-Piloted 

Simulation Model ••••••••••••••••••••••••••••••••••• 11 

Block Diagram and Example Time History of 

the Original Flight Path Command Logic ••••••••••••• 16 

Wind Waves at Sea •••••••••••••••••••••••••••••••••• 17 

U. S. Non-Aviation Ships ••••••••••••••••••••••••••• 19 

Axis Definitions for the Non-Piloted 

Simulation Model ••••••••••••••••••••••••••••••••••• 21 

Non-piloted Simulation Touchdown Velocity 

Results •••••••••••••••••••••••••••••••••••••••••••• 23 

Test Matrix for the Piloted Simulation ••••••••••••• 26 

Block Diagra~ of Vertical Velocity Damping 

Through Thrust as Implemented in the 

Simulation ••••••••••••••••••••••••••••••••••••••••• 27 

Throttle Control S~nsitivity ••••••••••••••••••••••• 29 

Head-up Display Format and Symboloqy 

D.efinitions ••••••••••••..•.••••••••••••.••••••.•.•• 31 

The Chair 6 Fixed-Base Simulator ••••••••••••••••••• 35 

Handling-Qualities Rating Scale •••••••••••••••••••• 37 

Mean Flight Times for the Piloted Simulation ••••••• 38 

Influence of Thrust-to-Weight Ratio, Engine 

and Airframe Dynamics on TouchdOlYn Sink Rate--

Baseline HUD with Velocity Command Control 

System ••••••••••••••••••••••••••••••••••••••••••••• 40 

Influence of Thrust-to-I-leight Ratio, Engine 

and Airframe Dynamics on Pilat Ratings--

Baseline HUD with Velocity Command Control 

~ystem·· ••••••••••• · •••••••••••••••••••••••••••••••• 41 

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LIST OF FIGURES_, continued 

Figure No. Title 

---'-

17. Comparison of Touchdown Sink Rate and pilot 

Rating as Influenced by Thrust-to-Weight Ratio 

for a Selected Condition--Baseline HUD with 

Velocity and Attitude Control systems •••••••••••••• 42 

18. Influence of Thrust-to-Weight Ratio, Engine and 

Airframe ~amics on Touchdown Sink Rate--

Augmented HUD with Velocity Command Control 

System ••••••••••••••••••••••••••••••••••••••••••••• 44 

19. Influence of Thrust-to-Weight Ratio, E~gine and 

Airframe Dynamics on Pilot Ratings--Augmented 

20. 

21. 

22. 

23. 

HUD with Velocity Control Syste·,n ••••••••••••••••••• 45· 

Comparison of Touchdown Sink Rate and pilot 

Rating as Influenced by. Sea State for a 

Selected Test Conditicn--Augmented HUD with 

Velocity Command Control System ••.•••••••••••••••••• 47· 

Influence of Thrust-to-weight.Ratio on the 

Positioning side Task--Augmented HUU with 

Attitude·. Command System •••••••••••• -•••••••••••••••• 48 

Distributi,c.n of Touchdown Sink Rate and Ship 

Deck Velocity--B~seline and Augmented HUD 

Breakdown for all Landings Completed for the 

Simula ticn ••••••••• , ................................ 48 

Influence of Thrust-to-Wei.ght Ratio on Thrustto-Weight Ratio Use Histograms--Baseline HOD 

with Velocity Command Control System ••••••••••••••• 50 

24. Influence of Thrust-to-~eight Ratio on Thrustto-Weight Ratio Use Histogra~s--Augmented HUD 

25. 

26. 

with Velocity Command Control System ••••••••••••••• 51 

Influence of Engine and Airframe Dynamics, and 

Thrust-to-Weight Ratio on Thrust-to-Weight 

Ratio Use Histograms--Augmented HUD with 

Velocity Command Control System •••••••••••••••••••• 53 

Influence of HUD Format and Control System on 

Thrust-to-weighf.: Use for a Selected Test 

Condi tion ••••• " • ••••••••••.••••••••••••••••••••••••• 53 

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LIST OF FIGURES, continued 

Figure No. 

27. Influence of Sea State on Thrust-to-Weight 

Ratio Use--Augmented HOD with velocity Command 

28. 

29. 

30. 

31. 

32. 

33. 

34. 

System ••••••••••••••••••••••••••••••••••••••••••••• 54 

Block Diagralll for the Modtfied Non-Piloted 

Simulation Model ••••••••••••••••••••••••••••••••••• 59 

Block Diagram of the Modified Flight Path 

Command Loqic •••••••••••••••••••••••••••••••••••••• 64 

Example Time History for the Modified Flight 

Path Command Logic ••••••••••••••••••••••••••••••••• 65 

Aircraft and Ship Motion Time History as 

Output from the Non-piloted Simulation ••••••••••••• 66 .. 

Aircraft and Ship Motion Ti~e History as 

Output from the Non-Piloted Simulation ••••••••••••• 67 

Comparison of the Influence of Thrust-toWeight Ratio on Touchdo~~ Sink Ratc--

Augmented ;;UD with Velocity Command System 

of Piloted Simulation and Non-Piloted 

Simulation ••••••••••••••••••••••••••••••••••••••••• 68 

Comparison of the Influence of Thrust-toWeight Ratio on Flight Time--Angmented HOD 

with Velocity and Attitude Command Systems 

of Piloted Simulation and the Original and 

Modified Flight path Command Logic of the 

Non-piloted Simulation ••••••••••••••••••••••••••••• 68 

35. Comparison of the Influence of ThLtlst-toWeight Ratio and Engine Dynamics on Tbuchdown 

Sink Rate--Augmentcd lIUD with Velocity 

Command System and Modified Flight Path 

Command System ••••••••••.•••••••••••••• ~ ••• ~ •••••••• 69 

36. Comparison of Thrust-to-Hcight Use Histograms 

as Obtained from the Piloted and Non-piloted 

37. 

Simulations •••••••••••••••••••••••••••••••••••••••• 72 

Variation of Thrust-to-wcight Ratio Use 

Histograms Due to a Statisticclly Small Number of 

Runs Used Per Da ta Pain t ••••••.••••.••• " ............. 73 

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LIST OF FIGURES, continued 

Figure No. Title 

A1. Block D.iagram of the A.I.L. Implementation in 

the Non-piloted Simulation ••••••••••••••••••••••••• A9 

A2. Examle Output from SHPREF.FOR with Plot of 

Ship Motion Specturm for the 00962 in Heave ••••••• A11 

A3. Example Output from SHPREF.FOR with Plot of 

Ship Motion Spe'ctrum for the 00962 in pitch ••••••• A12 

A4. Strip Chart Recording of Vertical Acceleration 

Due to Turbulence as Obtained from the AV-SA 

Fixed-Base Simulation Facilities and Used to 

Determine Magnitude of Turbulence Values for 

the Turbulence Modeling in the Non-piloted , .:~ . 

AS. 

AGoA7. 

AS. 

Simulation •••••••••••••••••••••••••••••••••••••••• A15 

Example Plot of Thrust-to-Weight Ratio Use as 

Output from the Non-Piloted simulation •••••••••••• A17 

Example plot Showing the Variation in Variance 

and- Means for Flight Time and Touchdown Velocity 

as Influenced by the Number of Runs Made in the 

N"n-piloted Simulation .•..••..• _ .••..•••.••.•..•.. • ~1B 

J:-... ample. pl"t Showing the variation in Variance 

and Means for Ship Motion as Influenced by the 

Number of Runs Made, Where Number of Runs is 

Seconds in This Case •••••••••••••••••••••••••••••• ~'9 

Example Output of Run Conditions and Results 

for the Non-Piloted Simulation •••••••••••••••••••• A20 

B1. HUD Hover Point Dynamics ••••••••••••••••••••••••••• B4 

82. Example Data Output from the Piloted 

Simulation ••••••••••••.•••••••••.•••••••••••••••••• B6 

83. Piloted Simulation Test Matrix ••••••••••••.•••••••• 810 

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LIST OF TABLES 

Title 

Example Lull Criteria ••••••••••••••••••••••••••••••• 3 

Simulation Environmental Conditions for the 

00963 Class·Destroyer •••••••••••••••••••••••••••••• 33 

Touchdown Velocity Statistics Indicating 

Number of 4andings with Sink Rates Greater 

Than 12 ft/s for Selected Test Conditions •••••••••• 49 

Influence of Thrust-to-Weight Ratio and 

Engine and Airframe Dyn'l-mics on' Fligh~ Path 

Command, Logic Sequence' Us'e ••••••••••••••••••••••••• 62 

Values for Variables Which Must ~ Edited into 

Program with Change in Sea State •••••••••••••••••• A16 

Fixed-Base Simulation Com~uter Variables arid 

Related Variables and Values ••••••••••••••••••••••• B2 

Variable Value3 Use in Correcting for the Bean 

Aerodynamic Force~ ••••••••••••••••••••••••••••••••• B3 

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;) 3ymbol Definition !' 

AC Aircraft transfer function 

, ~ Amplitude I 

AR Acceleration magnitudecos Cosine , 

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cmd Command I ,. 

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deg Degree 

ENG Engine transfer function 

• 

ELC Engine lag time I constant I 

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En Random phase angle 

("- ft Foot, feet 

g Gravitational constant 

h Altitude 

fls Significant wave height 

in Inches 

i(t) Ship motion component 

1), Pilot gain i..:.,. 

knts Nautical miles per hour 

m Meters 

msec Hillisecond 

n Random nl

1!':.ber 

s Laplace variable 

s Second 

sec Second 

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Symbol 

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Greek 

Symbol 

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LIST OF SYMBOLS, ccntinue'd 

Definition 

Time (seconds) 

Thrust 

Time increment 

Lag time constant 

Lead time constant 

Modal period 

Velocity 

Weight 

Position along the X axis 

True longitudinal acceleration 

Estimated true acceleration 

Estimate of high frequency acceleration 

Pilot transfer function 

Vertical velocity damping coefficient 

Increment 

Frequency 

Statistical variance 

Time cons.tant 

r~~ponerrt of the ship motion spectrum 

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LIST OF SYMBOLS, contin~ed 

Definition 

Roll angle (deg) 

pitch angle (deg) 

Yaw angle (deg) 

Ship heading (deg) 

Lap~ace variable 

Command 

Center of gravity 

Encounter 

Engine 

Landing pad 

Longitudinal 

Velocity 

Wind 

Wind over deck 

position along x axis 

Position along y axis 

Position along x axis 

Roll axis 

pitch axis 

Yaw axis 

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LIST OF SYMBOLS, continued 

Abbreviation Definition 

A.I.L. Approximate inverse Laplace transform 

HUD Head-up display 

N.A.S.A. National Aeronautics and Space Administration 

V/STOL Vertical/short takeoff and landing 

VTOL Vertical takeoff and landing 

A.C. A~titude command control system 

V.c. Translatiortal velocity command control system 

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I. INTRODUCTION 

The problem of landing V/STOL aircraft aboard destroyer class 

ships has been investigated in the past (References 1-3). Several 

meth.ods have been used to determine the feasibility of, and the 

control/disPlay systems needed, to ~ccomplish this task. 

Many of the researchers in this area began with the premise 

that for the success~ul completion of this· task, it would be 

necessary for the pilot/aircraft system to have the ca~abilit¥ ot 

in-phase chasing of the ship deck. The vertical task then was to 

start at a specified altitude, descend at a reasonable rate and 

begin to match the vertical motion of the ship deck. If the ship 

deck motion can be matched in both phase and amplitude, then it is 

only a matter of establishing a small relative descent rate and a 

reasonable landing ca~ be made. There are, of course, problems with 

this technique. In high sea states (5-6) the frequency of the ship 

motion is near the maximum piloting fr,equency (~4 rad/sec). In 

atte~pting to match both the phase and amplitude of the ship motion, 

the pilot is forced to operate at close to his break frequency and 

at fairly high gain. The aircraft must incorporate a high thrust--

to-weight ratio to achieve the maximum amplitudes in the time 

required. The combination of high piloting frequency and gain with 

high thr.ust-to-weight ratio can cause lags and a tendency to 

overshoot. If in addition there is a large system lag due to engine 

spool time, display lags, and pilot delay times, the phase lag can 

be.come excessive to the £)!)int of producing .large touchdown 

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velocities and an unstable system. This has been demonstrated in 

computer simulations (Reference 1). As a result it is generally 

concluded that deck chasing is not a reliable method for landing an 

aircraft with reasonable accuracies or consistently low touchdown 

velocities. This conclusion is confirmed in helicopter operations 

onto small ships. 

A second approach was based on the idea that the pilot could 

loiter until a lull in the ship motion occurred. Some of the 

research indicates that adequate lulls are not frequent occurrences 

or are too short in duration to· be useful. For example Reference 1, 

which investigated results for the two lull criteria given in Table 

1, determined that for the more conservative criterion no lulls 

occurred in the 00963 ship motion model over a period of 1800 

seconds. For the less stringent criterion, 52 lulls occurred in 

this time period, or the average of 1 every 33 seconds. This 

indicates lulls of very short duration. 

Reference 1 con~lU:d~~ that looking for lull conditions under. 

high sea states in order to make a lanqing is not very feasible. 

Other research conducted in the area of lull prediction (Reference 

2) indicates that lull conditions (defined a~ the time !rom which 

there are 2 successive peaks under the mean value of the positive 

peak amplitude envelope until 2 successive peaks over that value) 

occurs at the rate of 1 every 70 seconds and are of 10 to 60 seconds 

duration. Another consideration is found in Reference 3 where a 

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] Table 1: Example Lull Criteria 

Reference 1 LULL CRITERIA .. 

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• Motion .. Component Criteria Criteria 

Limit 1 2 

Longitudinal Vel. 2 ft/sec 3 ft/sec 

Lateral Velocity 2 ft/sec 3 fe/sec 

Vertical Velocity 2 ft/sec 3 ft/sec 

Pitch 1° 1.5° 

Roll 2° 3° 

Pitch Rate 2 o/sec' 8 a/sec 

Roll Rate 2 o/SI!C 8 a/sec 

No. of Occurrances 0 52 

Based on 1800 sec. of DD963 ship model motion for Sea State 5. 

description of sea trials performed using a SIl-2F helicopter 

indicated that pilots were often unable to determine visually wherr a 

lull was occurring. The above information indicates that in general 

it is not practical for fixed wing VTOL aircraft to loiter for the 

required time periods in the high fuel use state of hovering while 

waiting for optimum landing conditions. 

Another area of research involved the use of ship Motion 

prediction schemes. Research in this area has shown some promising 

results. Computer studies have shown that ship motion can be 

predicted with reasonable accuracies for 10-15 seconds' in advance 

(Reference 4). Given this capability, it has been demonstrated in 

computer studies using optimal control modeling techniques, that 

autopilot landings can be made with touchdown velocities on the 

order of 1 ft/sec in sea state 5 conditions (Reference 5). The 

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advantage to using a system in which the ship position is predicted 

in advance and updated as the approach progresses comes from being 

able to adopt a control strate9Y in which the aircraft is chasing a 

slower moving prediction point rather than thp. real time deck 

motion. This adds lead, thereby requiring less effort on the part 

of the system as demonstrated by adequate performance at thrust-toweight ratios of 1.05. The major problem here is that motion 

prediction for destroyer class ships in high seas h~s not been 

demons t~ated •. 

The research conducted for this report is directed at the 

question of what tht"us.t-to",:weight ratios, and vertical velocity 

damping are required to allow a pilot to make an acceptable landing 

given adequate situational information. 

It is clear that it is desirable to land with low touchdown 

velocities using low thrust-to-weight ratios, and without wasting 

time in the high fuel consumption state of a hovering loiter waiting 

for lulls in the ship motion. 

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II. ANALYSJ:S OF HOVER HEIGHT CONTROL 

TO provide some insight into the problem and the models being 

used in this research, the following facts are presented: 1) The 

current AV-8A Harrier aircraft has a maximum touchdown veloc.i ty 

limit of 12 ft/sec. 2) The 00963 ship motion mo1el has a maximum 

ship deck heave velocity (3 s~gma value) of 17.5 ft/sec. 3) In sea 

state 6, the ship has a heave velocity equal to or greater than 

12 ft/sec for less than 0.6' of the time. 4) The maximum heave 

veloci ty values occur in a region close to· the, ship motion mean 

position. 5) The lo~er velocities occur in the pe{l.k and trough 

regions of the motions. The above facts in conjunction with the 

ship motion histogram for sea state 6 would indicate that if the 

pilot did nothing but maintain a descent velocity of less than 

2 ft/sec, he would· touch down wi thin the AV-8A gear limits 

approximately 97% of the time. The sh1,p motion statistics show that 

a maximum heave velocity of 12.5 ft/sec is encountereu at the 2 

sigma heave amplitude for sea state 6 conditions. These same 

statistics show that the 2 sigma heave amplitude values are reached 

at reasonable time intervals of approximately 1 per minute. This 

frequency increases to 1 every 40 seconds for amplitudes of 1 foot 

bel~~ the 2 sigma value. These facts suggest a landing strategy 

differing from both the "deck chasing" and "lull waiting." This 

strategy has the pilot descend to the 2 sigma height above the mean 

deck position. ,At this height the pilot waits and watches the ship 

motion. If it cpp<!ars that a 3 sigma amplitude (high velocity) deck 

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motion is imminent, the pilot has enough height, and ther.efore lead 

time to begin an ascent, landing at the more desirable higher 

altitude. (lower ship and relative velocity) as the deck catches up 

to him. If it appears that the deck position is going to peak 

somewhat below his present altitude, there is again enough lead time 

to begin a slow descent and land near the crest (low ship velocity' 

position) of the deck motion. The 2 sigma height above mean deck 

position meets the desired criteria for the strategy. It offers an 

~asily obtained position without cont~nous deck chasing, provides 

the buffer needed to escape the high velocity portions of the deck 

motion, and presents the pilot with numerous landing opportunities 

f: 

without a lertgthy loiter period. To accomplish this, the pilot must 

be presented with a suitable indication of mean deck position and 

the 2 and 3 sigma values of deck position relative to the, aircraft's 

landing gear. This information requires measurement of the ship 

motion for several minutes. prior to the arrival of the aircraft and 

the transmission of this information to the aircraft for use in the 

head-up display. Real time information is also required to ShO'.o1 the 

pilot where the deck is currently positioned within the bounds of 

the probable travel. The pilot can then monitor the ship motion,. 

obtain an accurate deck position relative to the mean, and make his 

prediction as to how fast it is moving and where its position will 

be in a couple of seconds. 

An investigation of the proposed landing strategy was cond,.cted 

along two lines: namely, a non-piloted simulation and a piloted 

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zc;:;a. Pil. '. : :: = =::" , ;Pl,> ., ;;: !ilI%, ",_Ok '1$4~~ 

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simulation. The non-piloted simulation incorporates a linear pilot 

,model and vertical axis aircraft model. In addition there is a 

flight· path command logic section, a command flight path subroutine, 

a ship motion Subroutine, and a turbulence modeling subroutine. 

Input variables are' pilot gain, maximum thrust-to-weight ratio, sea 

state, vertical velocity damping coefficient, pilot time delay, and 

engine lag. 

The pllo.ted simulation W?ed, the chair 6 simulator at the 

N.A.S.A. Ames Research Center. This fixed-base simulator consists 

of a cab containing the normal., cockpi t controls and a single forward 

looking window th~ou9h which a visual image of the outside 

environment can be obtained. The visual image is provided by a 

camera-terrain' board imagin~ system. A head-up display can be 

superimposed on the outside scene to provide flight situat,ion 

information to the pilot. The cockpit also contains a unique 

throttle/nozzle contro~ quadrant used for V/~TOL simulation. 

Slightly modified versions of an existing math model of, the AV-8A 

(Reference 6) and an existing head-up display format were used'. The 

math model was linearized in the vertical axis and the display 

format simplified to represent only information needed to fly the 

ve:-tical axis. 

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A. CONTROL TASK 

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The task to be flown begins with the aircraft at an initial 

altitude of 45 ft above the mean deck position in a stabilized hover 

directly over the bull's-eye on the ship deck landing pad. The 

pilot uses the throttle to control altitude and vertical velocity to 

descend to all initial hover altitude. This initial hover altitude 

is the 2 sig~ value of ship deck position above the mean, as 

designated by a line on the head-up display (HUD). The pilot then 

lands at his discretion, h,ased on the ship deck motion information 

presented on the HtJD.. 

There are two vad.ations of this task. In the first variation, 

the ship deck motion bouncaries and reference lines are not 

displayed on th'P. mm and the pilot makes the landing withou.t the 2 

sigma reference. In the second variation, an attitude command 

control system replaces the translational velocity command control 

system which creates an effective sidetask in that the pilot must 

actively maintain the aircraft position over the bull's-eye, using 

the control stick, while; performing the vertical task. 

B. THE NON-PILOTED SIMULA~ION 

The non-piloted simulation was run through a control program 

which links the main progrilnT \'1i th flight path, ship motion, and 

turbulence subroutinp.s, data files, and subroutines for output of 

statistical and plotted data. Figura 1 shows the flow path' diagram 

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VASCO:t.FOR VASCO:t.DAT VASOUT.DAT I 

STAT. OAT CENSK.DAT I 

Vertical Input data Output Ou,tput Input data 

Axis I for initial data data for for ship 

Simulation conditions storage S tatht 1cal motion 

Control ... prog,ram Bubrout1ne 

progra", storage .. 

VPAlC. FOR ~ CENSM.FOR TRANO.FOR 

~ Hain Program Cenerates C eooral Ship Turbulence 

hover alt- Mot 10n. the' md ,Noise 

Flight path !tude com- ship motion generation 

ID command logie, ",and sub- generation subroutine 

pilot model, routine ,ubrout ino 

aircraft model, . 

st3tistics and IA plotting data J 

output 

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STAT .'FOR E!:Q.U !.!£!.2. !!:Q.U. 

Formatting A/C and TI'rI ratio Statistical 

and output ship motion as a per- variation 

of statistics ti ... c history cent of with nUmb~r 

time of runs 

Tabular Printout Plots Plota Plots 

Figure 1: Non-Piloted Simulation Computer Program Flow P~th 

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between the various programs and files. The main' program contains 

the flight path command logic, pilot neuromuscular model, and 

aircraft models. A block diagram showing the model trnnsfer 

. functions is shown in Figure 2. 

The rationale for developi.ng a non-piloted simulation was that 

it provided a relatively low cost test of the feasibility of the 

proposed landing strategy and its ability to l1Iinimize the required 

T/W. In.additiqn it was conjectured that comparison o~ the results 

of such a simulation with the piloted simulation would provide a 

test of the validity. of the intuitive O'l~dons underlying the assumed 

way the pilot would implement the strategy. perhaps not too 

surprisingly, the task turned out to be much more invloved than 

originall}, thought, and refinements to incorporate additional 

features suggested by the piloted simulation was a continous 

process. Nonetheless, the insight gained by the exercise was 

invaluable. 

The APPx:02'<imate Inverse Laplace Transform (A.I.L.) method (see 

Appendix A,), was used to solve the differential equations describing 

the pilot and aircraft transfer functions. From a review of the 

literature and looking at the task to be flown, it appeared from the 

beginning that the pilot transfer function would be the major 

problem. For the type of problem being looked at, a generic 

aircraft transfer function could be used, butbecau~e of the 

amplitudes and frequencies involved in the ship motion, lt was 

apparent that the time c6nstants used in the pilot transfer function 

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PILOT MODEL 

FLIGHT 

PATH 

LOGIC 

I- IPILOT 

DYNAMICS 

S~ -"H K, (TJ. • + 1) + _ e (TG s + 1) 

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NOTION 

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SEA STATE CON~ITIONS 

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AIRCRAFT VERTICAL AXIS MODEL 

POWER 

gravity PLANT 

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ELC s·+ 1 s2 - Zws 

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SEA STATE CONDITIONS 

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would be more critical. It was. also desired to keep the pilot 

transfer function separated from the aircraft transfer function as 

much as possible, so that either the pilot or aircraft portion of 

the program could be moved as a block, either for use in other 

programs or to faci1i tate looking at other systems in the curr.ent 

program. This need to separate the transfer functions complicated 

the A.I.L. equation set-up. 

C. ANALYTICAL MODELS. 

1. Aircraft Model 

Th.e dynamics of the aircraft vertical axis are represented by 

two first order cascaded tra.nsfer functions (Figure 2), one 

representing. the airframe vertical velocity response to a thrust 

change, and the second representing the engine response to a power 

lever input. The feedback from the aircraft to the pilot is assumed 

to be aircraft altitude only. 

Only the vertical axis of th~ aircraft is modeled. The first 

order powerplant transfer function representing powerplant lags is 

dT/l'l 1 

dT/W = (ELC s + ,) c 

Cascaded to this is the aircraft transfer function: 

h g 

dT/W = (52 + z 

w s) 

To give a total vertical axi's transfer function of 

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h (g/ELC) 

dT/WC = [93 + (l/ELC + ZW)S2 + (Zw/ELC)s1 

(3) 

A limiter iR put on the pilot output of T/\i to prevent any input 

above the maximum T/W ratio or below 0, i.e., the pilot can not 

command a negative thrust. 

The engine lag constant, ELC, is one of the parameters on which 

the control requirements depend. 

2. pilot Model 

The pilot transfer function was first set up us.ing a first 

order lead-lag multiplied by the Pade approximation (Figure 2). 

This transfer function was looked at using the Linear Systems 

Analysis Program (LSAP). Bode plots, root locus plots, and time 

history plots were looked at to determine values for lead and lag 

time constants which gave ~he best results for both a step input and 

a sine-wave input. The values which subjectively produced outputs 

similiar to those expected from a piloted simulation were then used 

in the non-piloted simulation. Because the computer program is 

nonlinear, several runs were made to determine which values for lead 

and lag still gave a good combination of rise time for step inputs 

and small phase and position errors in following a sine-wave. 

Values for the lag time constant, TG, of 0.1 sees and for the lead 

time constant, TL, of 0.5 sees were finally selected. These values 

are within the range of values usually quoted for pilot models 

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(Reference 7). 

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Because a pure time delay could be programmed into the nonpiloted simulation, the Pad~ approximation, which was needed in the 

analysis using LSAP, was replaced. 

The pilot transfer function is 

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p 

= K (TL s + 1) 

P (TG s +1) 

The pilot's logic is represented by a logic section in the 

(4 ) 

simulation. The logic section consists of 4 basic sections. The 

first section is a series of logic statements which determine which 

of the other 3 sections will be used to provide the commanded flight 

path. These sections will be referred to subsequently as ABORT ro 

HOVER HEIGHT, CHASE, and RUN FROM. 

If the aircraft is more than 6 ft above the ship deck, or has 

followed the ship deck be.low a specified abort chase altitude, or 

has exceeded a specified vertical velocity (a function of the 

maximum available T/\"), the flight path' command logic enters the 

ABORT TO HOVER HEIGHT section. The hover height is the 2 sigma 

value of the ship deck heave above the ship deck mean position. The 

CI~SE sequence is entered if the ship deck is within 6 ft of the 

aircraft and the ship deck velocity is less than 2 ft/sec 

(approaching the aircraft) and decreasing. When the CHASE sequence 

is entered, the aircraft flight path is commanded to ma,tch the ship 

motion. The RUN FROM sequence is entered if the ship deck velo.city 

is greater than 2 ft/sec or if the velocity is increasing; i.e., the, 

ship is accelerating toward the aircraft. In the RUN FROM sequence, 

the commanded altitude is the ship position plus an exponential 

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smoothing function (a function of time, based on when the sequence 

.. was entered}. Because the exponential function dies out with time, /'. 

the aircraft is prevented from climbing to excessive altit-,udes 

before one of the other sequences is initiated. A block diagram and 

example time history of the flight path commana logic prior to thepiloted simulation are presented in Figure 3. 

After results were obtained from the piloted simulation, models 

representing the time the pilot spends flying a part~cular portion 

of the task and delays in perception were added. Since the pilot 

scans the situation and instruments and corrects errors in a 

sequence rather ,than in parallel, the time delay was divided into a 

combination of the pure delay and a gap where'the input was 

maintained at a given value for the time a pilot could be considered 

flying another axis or scanning other instruments and therefore not 

activily flying the vertical task. The length of the time delay, 

the length of the gap, and how often the gap occurs are variables 

that can be initialized at the beginning of a r~n. In addition 

there are inputs for pilot preception error noise and pilot internal 

noise. These are discussed in more detail later. 

3. Ship Model 

An understanding of sea state can be gained by referring to the 

chart in Figure 4. Sea state is shown with the associated "rind, 

wave heights, lengths, and periods'. Sea state 6 is considered , . 

significant in that it is estimated that operational capability 

under these conditions would provide for use of aircraft 67% of the 

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lYES ~ i R;liAINOER OF IS I/O IS tlO IS NO IS NO IS no I ''''''''L - '-t FlAC] H "'O.J-, "U-' '---0 ' ---c ZO r-o SSLS .LT. - r -1.0. .LT. .LT. .LT. 2.0 .AND. 

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Height 

Ship 

Deck 

Hean 

I' ZWOT 0.0 

YES ! YES TES YES 

SAl, SAl, SAJ are generated by 

the hover altitude subroutine, SET . 

1N4. FOR. SSL1, SSL2, SSL3 are FLAG3 ,- generated by the ship motion -0.0 J generating subroutine, GEl/SH.FOR. SI - SAl I 51 - SSLl SI - SSLI + 

S2 - SA2 S2 - S5L2 EXP(I.9*(1-!E/2»-1 

51 is the cO~~3nd input to the S3 - SA3 53 • SSL3 52 - SSL2 

pilot transfer function. SET S3 - 5SL3 

FLAG3 I' I 

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Aircraft follo~s flight path 2) Aircraft ~ithin 6 ft. Aircraft descends f; Rapidly rising con~and to hover altitude. 

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of 4escending ship deck, belo~ abort chase ship deck, ~irenter 'chase' mode., altitude, abort to 'enters 'run from' 

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Aircraft descent rate hover altitude. mode until, in 

d 1 5*ZWTDT b thia case, touch- excee 9. , a ort do~ occurs. to hover altitude. 00 

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Time 

• ··lloveiAltitude 

--Aircraft Flight Path 

-Ship Deck Position •• 'Commanded Flight Path 

Figure 3: Bloc~ Diagram and Example Time History of the Original Flight Path 

Command Logic 

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WIND WAVES AT SEA 

Correep~nding v31ues lie on a vertical line 

1 WIND VELOCITY (knte) 4 5 6 7 8 9 10 jo ~ 40 50 6b it_ 2 BEAUFORT WIND I 2 3 4 5 6 7 8 9 10 11 

and DESCRIPTION light light gentd moderate r.resh strong mod !cruh stt ~hol IItOrtl 

air breeze breeze breeze breeze bre.:!ze rat;' gale ong IIle 

gale gale 

--- 3 Rt:QU I RED rETClI retch Is the number of miles a given I 2Jo 30

1

(mi) wind has been blowing over open lIater. 5,0 10,0 T .1 0 4~0 ~ !lro 6~0 7~0 -r-- 4 Rr:QUIRCn 

DURATION 

WIND 

(hr) been 

~lration 

blowing 

is the 

over time open a water. 

given wind has 51 

~I 2~ 2'5 3~ ", -.- -f ,,,2l __ 5 WAve HEIGIIT , 21 4 (Cr~st to Trough (ft» 1 -, wh'ite ~ ~ 10 ~ 2~ 2'5 ~O j 1 .1 J. 40 ~O ~O 

6 'Sr.A STATt 1 2 r§F~ 4 5 6 7 8 --- and DESCRIPTION smooth sl1~ht moder ough very high very percipitoul .... 

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ate rough high - r --y -~{ 7 WAVE pr.RIOn (sec) 1 2 3 4 6 8 10 12 14 1 18 2 ' 

8 \lAVE LENGTI! (ft) 20 40 ~O, 80 ioo 150 ~00 300 4¢0 ~O 600 8PO IPOO~~o:. ~~~_ ') lIAVF. VELOCITY (knt s) ~ 10 1> 2P 2~ 3P 3~ 4b 4~ J2..51 ~ 10 PARTICLE VELOCITY (ft/sec) 1 2 ~ 4 $ 6 j 10 11 14 

11 WIND VF.LOCTTY (knts) .1.

4 5 6 ! 8 9 1.0 20 30 ~ 50 6p "ir--- , -'----

This table applies to wsves generated by the local wind and dows not apply to awell orginating elaewhere. 

WARNING NOTE: Presence of swell maken accurate wave observations exceedingly difficult. 

a. The hel&ht of waven 1s arbitrarily chooscn as the height of the higheot 1/3 of the 

W6ves" Occasional waves by interference between wavea or between vavel and Iliell 

may be conSiderably larger. 

b. The above values lire only approximate due to lack of precise data and to the difriculty in expressing it in a single easy way. c. Below the surCace the wave motion decreases by 1/2 for every 1/9 of • wave length of depth increase. 

Figure 4: Wind Waves at 3ea 

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time, in the North Atlantic during January. If operations can be 

conducted only under sea state 3 conditions, operations would be / 

limi ted to 3" of the time. 

It should be noted that ship motion may be considerably greater 

than indicated by the wave amplitude values. For example in sea 

I 

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i 

state 6 the wave height rraxililum is 20 ft from crest to trough. The 

heave motion for the 00963 class destroyer can approach 40 ft from 

1 

! } I ~-' -" crest to trough under sea state 6 conditions. 

! "\ ( ---- I I I 

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The OD963 Spruance class destroye~ model was used for both the 

piloted and non-piloted simulation. This ship' was chosen because a 

motion program for use with the piloted simulation was already 

I 

i available. The 00963 is considered to be typical of the' type of ship 

f 

t 

I 

from which VTOL operations could be conducted. The non-piloted 

simulation was set up so that data for other ships, as contained in 

Reference 8, could also be used. To provide some perspective of the 

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ship used, Figure 5 shows a listing of ships according to type and 

class, information'on the number of helicopters (assumed replaceable 

by VTOL aircraft' that can be carried, dimensional information, and 

a chart showing displacement. . A single degree of freedom ship model was programmed using the I 

method outlined in Reference 9. The data for the 00963 class 

I. ' I~ 

destroyer ~ere obtained from Reference 8. The model consists of the 

--' 

superposition of twelve sine waves, six representing the heave 

motion at tile ships center of gravity and six representing the pitch 

motion. The pitch motion is mUltiplied by the appropriate moment 

18 
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... 

<P 

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SIIIP TYPE CLASS IIELI- DIMENSIONS DISPLACEMENT COPTER T . • • - RELIANCE C. C. CUTTER ~'HEC 615 1 210.5 34 10.S 

BEAR C. C. CUTTER WHEe 901 1 270 36 13.S - GARCIA FRIGATE 'I'F 1040 * 1 U4 44 24 . WEST WINO ICE BREAKER WAGS 83 2 269 63.S 29 . O.H. PERRY FRIGATE FFG 7 * 2 445 45 24.5 - ". 

KNOX FRIGATE FF 1052 * 1 436 46.8 24.8 - IiACKINAW ICE BREAKER WAGS 83 1 290 74 19 . , 

SPRUANCE DESTROYER 00 963 ** 2 563.2 55.1 29 - BELKNAP G. mS5LE CRUISER CC 26 1 547 54.8 26.8 - , 

GLACIER ICE BREAKER WAGB 2 309.6 74 29 . I 

, 

AUSTIN AY.PllIB. TRANS. DOCK ·LPD 4 6 570 100 23 . I 

liARS CO~[;JAT STORES AFS 1 2 581 79 24 . 1110 JlliA AHPIIIBIOUS ASSUALT LPH 2 *** 11-20 602 84 26 . KILAUEA AMHUlHrrON T-AE 26 2 564 81 28 . WICHITA REPLEllISIlMENT OILER AOR 1 2 659 96 33.3 

BLUE 

SACRAI'lElITO 

RIDGE AI'~HIBIOUS 

FAST COMBAT 

COMMAND 

SUPPORT 

Lee 

AOE 

19 

1 2 

1 620 

793 107 

82 29 

39.3 - SAl PAll Al'lPHIBIOUS ASSUALT LHA 2 19-26 820 106 26 . --- AMl'IlIBIOUS ASSUALT LHD 1 19-26 840 106 26 . 

L II II 10 20 30 40 50 

" On:ER SHIPS CONTAINED IN REF. 4 DATA BASE 

** SHIP HCDELED IN THIS SIMULATION (ft) (ft) (ft) TONS x 1000 

*** THIS snu' IS ALSO HODELED AT N .. \.S.II. MlES ~-- ----- - ---

figure 5: U. S. Non-Aviation Ships 

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arm and the vertical component is extracted and added to the heave 

motion to obtain the heave motion at the landing pad. Reference 9 

reports that this method gives ship motion accuracies to within 5~ 

of a model containing 30 superimposed sine waves per axis. 

This ship motion program uses the Bretschneider wave spectrum, 

transformed based on ship velocity and heading relative to the 

waves. The ship motion spectrum is then obtain~d by combining the 

wave spectrum with the ship response functions and phase 

differences, ~ii' and the six amplitudes, ~, and frequencies, wen' 

are extracted. The appendix contains more information on how this 

is accomplished. A component of the ship motion is then represented 

by 

6 

i(t) = r A cos(Went - ~ii + En) 

n=l n· 

(5) 

The phase difference~ are directly available from the ship motion 

data base information contained in Reference 8. En is a random 

phase angle which is obtained from a random number generator with an 

output scaled to give valu~d between 0 an 6.242 radians. Figure 6 

shows a graphical definition of the axis system used in the 

simulation. 

4. Turbulence Model 

The turbulence model (Reference 9) consists of white noise, 

shaped by the following filters: 

20 

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tlesn Deck 

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Position 0 

. - - - - - - Pr7-'=t=':i:::j:=-----...... --.." 

NON-PILOTED S IlruLATIO~1 

AXIS DEFINITION 

~ ·::S' , 

North 

t Ship heading 90. 

Station-keeping 

Point 

NON-PILOTED SI}nTLATION SHIP 

OREIt:tATION TO WAVES 

Figure &: Axis Definitions for the Non-piloted Simulation Model 

--- -=.~-----:.-,.-~~-,.~--

, 

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---

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t. C,.tE" .4.' '4- .. 

(6) 

Where ~ is turbulence induced accelerations on the aircraft in 

g'8. The bandwidt~, wn' was obtained from RefeLence 9. Values of 

the variance, Gv ' were obtained from a strip-chart recording of the 

turbulence induced vertical acceleration for the AV-SA during a 

fixed-based simulation. The term containing the simulation time 

increment, TC, is a correction to the power spectr~~ to allow for 

digitization. The quantity, n, is a random number from a Guasian 

distributed sequence with zero mean and unity variance. 

o. RESULTS FOR NON-PILOTED SIMULATION 

Touchdown velocities obtained from the non-piloted simulation 

incorporating the flight path command logic devised prior to the 

piloted simulation are shown in Figure 7. This figure shows the 

average over many runs during which damping constants, and various 

parameters in the flight path command logic, as well as pilot gain, 

were being manipulated to roughly determine the range of values that 

could be expected for touchdown velocities. These runs were also 

monitored as they were occurring to determine (subjectively) which 

values provided realistic piloting responses. As such, the results 

should be interpreted as a rough indication of the trends. 

22 

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0 -0.1 

Y:: 0 -0.4 

'("Oil: - 0.7 

MEAN 

TOUCH-OCM!f 

VELOCITY 

(ft/see) 

20.0 

15.0 

10.0 

5.0 

0.0 

1.0 

SEA STAT! 6 CONDITIONS 

~ c· 

1.05 

TBRUST-TO-IIl:IC)tT RATIO. T/W 

. 

1.10 

Figure 7: Non-piloted Simulation Touchdown Velocity Results 

Figure 7 shows that in sea state· 6 conditions, the orginal 

flight path command logic suggested the very surprising res~lt that 

the landing task could be accomplished without ever exceeding a 

touchdown velocity of 12 ft/sec, for T/W values.~o~ to 1.01 and 

with engine lags as high as 0.7 sec. Moreover, the mean value of 

touchdown velocity was only mildly dependent on T/W and engine lag, 

ranging between 5.5 ft/sec and 6.5 ft/sec. Flight times from the 

initial altitude of 20 ft above the mean deck height were again only 

mildly dependent on Tl\'l and engine lag, ranging from 20 to 30 sec. 

The unexpected nature of the preliminary non-piloted simulation 

results strongly indicated the need. for experime:ntal ver.ification. 

This experiment was performed on a fi~ed-base simulator at N.A.S.A. 

23 

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Ames Research Center using an existing model of the AV-8A. It was 

recogni~ed from the outset that an important requirement of this 

experiment was the need to supply the pilot with exactly the same 

n 

information assumed in the construction of the command logic in the 

non-piloted simulation. Fortunately, an existing head-up display 

format termed SUPAR HOO (Reference 10) was available and wa~ 

modified by removing all augmented information and adding the basic 

situation information of mean deck position, and the 2 sigma and 3 

sigma deck positions relative to the mean. 

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III. SIMULATION EXPERIMENT 

A description of the· test matrix, aircraft and ship models, 

simulation cab, and results obtained for the piloted simulation are 

presented in the following sections. 

A. TEST MATRIX 

Based on the information gained from the non-piloted 

simulation, the test matrix shown in Figure 8 was set up. Sea 

state, thrust-to-weight ratio, vertical velocity damping through 

thrust, pilot, and HUD format were the variables. The task was then 

simulated using the fixed-base simulation facilties as further 

descibed below. 

B. AIRCRAFT MODEL 

1. Basic Aircraft 

The AV-8A model used in ~~e piloted simulation is described in 

Reference 6. It is a real time digital computer program developed 

to simulate the take-off and landing of V/STOL aircraft aboard 

ship. The unmodl.fied aircraft model includes nonlinear 

aerodynan.ics, engine and reaction control systems, stability 

augmentation and actuator dynamics, and a landing gear modeL. 

25 

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lIUD ItUDl IIUD3 

FORtlAT VEL. COHo SY~. VEL. COM. SYS. 

~ tNCIN!': 0.3 0.7 0.3 0.7 

Sr.A lAG 

STATE ~ ...... ,.., U'\ " 0 ..... C""\ V'\ ,.... 0 ...... C"'1 "',.... 0 .... M 11\ ,..., 0 

~~c:C!-: o 0 0 0 .... C! q C! C! ~ 0600 .... . . . . . . . . . . ...... ...... ...c ......... ..... ..... ..... .... ...... ..... .... ...... r"'4 .... ..... ..... .... ...t ~ 

. -0.0 o 0 0 0 0 0 o 0 0 0 

0 • -0.2 o 0 o 0 0 0 0 0 0 0 

-0.4 . o 0 0 0 0 0 O. 0 0 0 

-0.0 0 0 0 0 0 0 0 0 0 0 

4 -0.2 0 0 0 0 0 0 0 0 0 0 

-0.4 0 0 0 0 0 0 0 0 0 0 

-0.0 0 0 0 0 0 0 0 0 0 0 o 0 0 0 0 0 o 0 0 0 

6 -0.2 0 0 0 0 0 0 ·C 0 0 0 0 0 0 0 0 0 0 0 0 0 

-0,1, 0 0 0 01 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 O. --

o IIIDICATES AREAS OF IIATRIX TESTED 

Figure 8: Test Matrix for the Piloted Simulation 

/ 

HUD3 

ATT. COM. SYS. 

0.3 0.7 

.... ,.., "" ,... 0 ...... ,.., "'. ,.... 0 

~ ~ ~ ~ .-: o 0 00 .... . .... .... ................. .... -:-t ..... .... .... 

0 0 0 0 0 0 

0 0 0 " 0 () 

0 0 0 0 0 0 

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2. Modifications to the Simulation Hodel 

only the turbulence portion of the airwake simulation was 

used. !n order to create a simulation environment. as close as 

possible to that assumed in the non-piloted simulation, the mean 

wind and its variation with height above deck were not used. In 

addition, the ground effects inherent in the AV-SA were deleted from 

the simulation to match the non-piloted simulation, and to more 

effectively isoVlte the parameters being studied. 

A block diagram of the vertical velocity damping through thrust 

is presented in Figure 9, along with a definition of the variables 

used. 

_6!-IOOcr_-i~1 AIRCRAFT TRA..'1SFER 1-----.--1> h 

++ FUNCTIOI. 

CRDa! - Ale CAIN 

CHDOT - POYER L-~tR ~"CLt (deg)/(ft/sec) 

Ale CAIN - 0.45 (ft/sec2)/POWER LEVER &"CLt (des) 

Figure 9: Block Diagram of Vertical Velocity Damping' 

through Thrust as Implemented in the Simulations 

27 
,.' 

~ ... 

3. control AUgmentation System 

/ 

,/- - In running the simulation, the throttle quadrant was used with .. a fixed range of travel throughout the test matrix. As a result, 

the throttle sensitivity (vertical acceleration per degree of 

throttle) varied with thrust-to-weight ratio. A plot of commanded g 

per degree of throttle travel may be found in Figure 10. 

Two control systems were used for the fixed-based simulation. 

For the portion of the simulation in which only the ,vertical axis 

was being flown, a translational velocity command system was used. 

With this system the stick controls velocity in the longitudinal and 

, r-' . lateral directions. When the pilot leaves the stick centered, the 

aircraft maintains position (zero velocity). The pilot only had to 

control the throttle to perform the task. 

The second control system was an attitude command system. It 

was used to provide a positioning side task for the pilot. Using 

this system, the pilot flew the vertical task using throttle and had 

to actively mai:ntain position"over the ship deck by commanding 

attitude through the stick • 

. ~, 

28 

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0.0160 r--'-"""r-~-r---'--r--'-_T---'r--..., 

VERTICAL 

ACCELERATION 

PER DECREE 

THROTTU: 

:J.OlS0 

0.0140 ..... --'''--. ....... _--L._-'-_ ....... _~ __ --''-...... _ .... 

1.09 1.0S 

THRU:>r-TO-Il.EICHT RATIO, TN 

Figure 10: Throttle Control Sensitivity 

1.10 

I 

The transfer functions fOf the control systems are as follows: 

Translational velocity command. system: 

v 

x 

1.25 

s + 1.25 

= 7 f~/s l.n 

Attitude command system: 

-i...= 4 

~cmd (s2 + 3s + 4) 

4>cmd --- 11°/in 

4>lat stick -

v = __ 1._2,....5~ 

Y s + 1.25 

7 f~/s 

l.n 

e 4 e-= 

cmd (s2 + 3s + 4) 

o long stick 

(7) 

(8) 

More information on the control Systems being used in fixed-based 

simulations may be found in Reference 11. 

29 

i 

" I 

I 
i, . ( 

4. Head-Up Display 

Two HOD displays were used in the piloted simulation. These 

are referred to as HUD1, and HUD3, and are shown in Figure 11. HUD3 

was used most extensively in the testing. HOD3 contains the ship. 

position reference lines and aircraft hover height and "abort chase" 

altitude lines. HOD 1 contains only the symbol indicating current 

ship deck position. HUDl was used to determine how much (if any) 

advantage there was to displaying the extra information on ship 

motion boundaries. It should be noted that the HUD3 format 

collapses into the HOD 1 format for calm seas (sea state 0). 

The head-up display superimposes vertical and horizontal 

situation information. The trident symbol is fixed in the center of 

the display and shows the aircraft's vertical position as the 

distance of the three 'pads' of the trident fro~ the top of the ship 

deck reference symbol. In addition, the trident provides horizontal 

situation indications to the pilot when actively flying the sidetask 

with the attitude command control system. When the sidetask is 

flown, a dagger is added. to the display to indicate the undisturbed 

position of the deck bull's-eye. The distance between the trident 

and the dagger shows the error in lateral and longitudinal 

position. The HOD format also contains symbols showing a horizon 

line, pitch bars, pitch reference, and side slip. The lines added 

for the experiment show the 3 sigma, 2 sigma, abort chase height, 

and mean deck positions. 

,~-

30 

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cr--~t----t--_ PITCH REFER~CE SYMBOL 

A.~ SIDE SLIP BALL 

PITCll LADDER 

k-.:::::-r_-,.._....,·-+ ___ RORIZON/DlRECTIONAL 

RIBBON . 

AIRCRAFT POSITION A.~D 

OREINTATIO:l REFERENCE 

SYMBOL 

SHIP DECK REFE~CE 

SYMBOL 

.!ill!ll. 

POSITION OF DECK BULLS EYE 

USED FOR POSITIONING SIDETASK WITH ATTITUDE COMMA.~ 

SYSTEM 

3 a DECK MOTION BOUllDARIES 

_____ -t---2 a ROVER HEIGHT REFER~CE 

_

-=-.:.-.:.-:;,:-;.,;-:,.::-:.::===-:::::--t---ABORT OIASE REFERENCE LI:m e::::::J 

MEAN DECK POSITIO!f 

Figure '1: Head-up Display Format and Syr.lbology 

Definitions 

31 

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The equations describing the transfer function for moving the 

dagger relative to the trident are described· in appendix B. To 

assist the pilot in performing the station keeping task the dagger 

motion is augmented with input of aircraft true velocity, the 

estimated a~celeration, and stick position. The HOD symbology was 

driven by a Digital Corperation PDP-1VES computer at an update frame 

time of 110 msec. 

C. SHIP HODEL 

Ship dynamics are'modeled as six degree of freedom sinusodial 

I ~ . motion. The ship is' assumed to have a fixed mean position about 

which it oscillates. Wind over the deck is composed of a steady 

induced wind equal to the ship speed plus a separate North or East 

component of natural wind which can be specified. At present a 

turbulence model developed for the 00963 class destroyer is used 

(Reference 8). This subroutine calculates the free air turbulence 

as well as ship wake turbulence. The latte~ varies with position 

relative to the landing pad. Table 2 gives an indication of the 

environmental conditions for the simulation. 

D. SIMULATION FACILITIES 

,. Simulator 

The fixed-base chair (Ch. 06), is used primarily to develop 

controls and head-up displays for use in VTOL aircr~ft and 

32 

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Table 2: Simulation Environmental Condition~ for the 00963 Class Destroyer 

CONDITION SEA V \l 1/1 1/Iwod V vwod 11 s s w w 8 

:;UHBER STATE (knts) (deg) (deg) (deg} (knts) lknts) (m) 

1 6 25 120 -60 -30 25.00 43.30 5.49 

7 4 25 105 -75 -10 17.68 34.15 2.10 

14 0 10 - -68.6 -30 8.07 15.00 0 

0,. Os 01/1 ° CJ ° CJ 

x Y ° eg cg 

Z 

cg xl Yip 

(deg) (deg) (deg) (m) (m) (m) (m) (m) 

1 3.13 1.05 0.45 0.24 0.71 1.51 0.45 0.63 

7 1 11 0.34 0.17 0.05 0.27 0.60 n.12 0.18 

14 0 0 0 0 0 0 0 0 

o· o· o· o· o' (I' I" o· ~ a IjI x Yeg z eg eg XlI' Ylp 

(deg/sec) (deg/sec) (deg/sec) (m/sec) (m/sec) (m/sec) (m/sec) (m/sec) 

1 2.00 0.90 0.36 0.15 0.41 1.10 0.32 0.46 

7 0.88 o.n 0.18 0.04 0.21 0.53 0.10 0.20 

14 0 0 0 0 0 0 0 0 

~ 

i.._ .. ____ . 

I 

T 

0 

(s,,:) 

15.13 

10.60 

-

° zlp 

(m) 

1.67 

0.65 

0 

o· 

Zip 

(m/ sec) 

1.31 

0.59 

0 

'0 

r=;3) 

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helicopters (Figure 12). Configured as a single-seat cockpit, it is 

equipped with a conventional stick and rudde~ pedals with adjustable 

trim. The chair is provided with a virtual image TV display, on a 

single forward looking window. The outside scene is provided by a 

camera-terrain board system. A head-up display can be superimposed 

on the outside scene to provide flight information to the pilot. 

The cockpit also contains a unique throttle/nozzle control quadrant 

used for V/STOL simulation. Aircraft dynamics are provided by a 

Xerox Sigma 9 digital computer. The aircraft dyna~ics are updated 

at a frame time of 55 msee for the AV-SA model. 

2. pilots 

Two pilots were u~ed for the simulation. Both pilots were used 

in the sea state 6 portion of the test matrix. Pilot B filled in 

.the remaining tested portions of the test matrix. 

pilot A is a research engineer at NASA Ames Research Cen~er. 

He has been flying both the fixed and moving base simulators fo,r 15 

years. Most of this time has been spent using the AV-SA and YAV-SB 

simulation models. He was also heavily involved in simulations of 

advanced lift-fan transport (ALFT) aircraft and the Navy research 

and technology aircraft (RTA). 

Pilot B is also a research engineer at NASA Ames Research 

Center. He has been flying both the fixed and moving base 

simulators for over 4 years. Host of his simulation time has also 

been with the AV-SA and YAV-SB models. He also has some helicopter 

simula tion time. 

34 

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r--.. Figure 12: 

a. 

b. 

c~·~~t..i.:;<:' ... ~ 

PE pOOh: 

i~~t~ ~,~~___ -...~~-: .. ~' ~~r: 

.~~i!m,~"<}Gl : :~. -.(.tt> ~""JEt'ilJ >-,,:': 

-:::;,~W'~ 

.. ~".; . . ,..-~~~ ,.-

outside View of the Cab 

Inside View of the Cab 

The Cha1r 6 Fixed-Dase Simulator 

35 

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--------------~.--------~--~------~ .•. ~-,---------------

The pilots were asked for comments and a pilot rating following 

each series of the five runs that made up a data point. The pilot 

rating is based on the Cooper-Harper rating scale (Figure 13). More 

information on the Cooper-Harper scale may be found in Reference 12. 

E. OPE~TIONAL TASK 

The task to be flo ... m was' the same as described in the nonpilot~ simulation. The aircraft was positioned at an initial 

altitude of 45 ft above the mean deck position, in a stabilized 

hover, directly over the bull's-eye on the ship deck la~ding pad. 

The pilot, through use of throttle, was then to .. control altitude aad 

vertical velocity in descending to an initial hover altitude. The 

intial hover altitude was the 2 sigma value of the ship deck 

position, above the ship deck position mean, as designated by a line 

on the HUn (HUD3). The' pilot was then to land at his descretion, 

aided by the ship deck motion information presented on the HUD. 

There were two variations of the task. In the first, the HUn 

was reconfigured so that the ship deck motion boundaries and 

position reference lines were not displayed (HUD1). The pilot was 

then to complete the task wi th only the deck position symbol as a 

reference. In the second variation, the control system was changed 

from the translational velocity command system to an attitude 

command system. This was done to provide the pilot with the 

sidetask of actiVely using the stick to maintain the aircraft 

position over the deck bull's-eye while performing the vertical 

task. 

t '. 

I 

! 

! 

I 

i 

I . i 

I - ", ~ 

. ! J.. 

~<j , ., 

I 

I 

! 

! 

I 

I 

I 

! 

./ 

/ 

,/ 

./ 

.< 

.'~ 

/ 

,' . 

1 -_ ...... 

... 'i 
t.I 

..,j 

) 

ADEQUACY FOR SELECTED TASK OR 

REQUIRED OPERATION-

) 

AIRCRAFT 

CHA RACTE R ISTICS 

Excellent 

Highly desirable 

r-----------------------~~'1 Good Negligible deficiencies 

Yes 

Deficiencies 

warrilnt 

improvement 

Deficiencies 

require 

improvement 

Fair-Some mildly 

unpleilsant deficiencies 

Minor but annoying 

deficiencies 

Moderately objectionable 

deficiencies 

Very objectionable but 

tolerable deficiencies 

Major deficiencies 

Major deficiencies 

Major deficiencir.s 

Major deficiencies 

DEMANDS ON THE PILOT IN SELECTED 

TASK OR REQUIRED OPERATIONPilot compensation not a factor for 

desired performance 

Pilot compensation not a factor for 

desired performance 

Minimal pilot compens.?tion required for 

desired performance 

Des;red performance requires moderate 

pilot compensation 

Adequate performance requires 

considerable pilot compensation 

Adequate performance requires extensive 

pilot compensation 

Adequate performance not attainable with 

maximum tolerable pilot compensation. 

Controllability not in question. 

Considerable pilot compensation is required 

for control 

Intense pilot compensation is required to 

ret~in control 

Control will be lost during some portion of 

required operation 

PILOT 

RATING 

Cooper·Harper Ref~NASA TNO·5153 

°Definition of required operation involves designation of flight phue and/or 

subphases with accompanying condi!ions. 

Figure 13: Handl1r.g-Qual1ties Rating Scale 

~~~$;~~~li';:itw: ;i;l~'~ ~~~~~;:~~;~'~;~:;:-~~~j--:2L~~~~n;;: .t~~~:~'; f~~-~~:=='~~ ~_~~~--=== _, _ ____ boo 

C? .' ;., 

,.. 

-.. 

.) 

-. 

\. ----".~ .. ..--.-

'ii1\.' .~ . .:... 

'f \ \ .......... 

............. '''''''' 

...... 

,;t:o 

ror. 

.. ~ 
"- '" ~. 

------------------_._--- -- ----'---

F. RES~TS FOR THE PILOTED SIMULATION 

1. Task Performance and pilot Ratings 

A record of the test matrix and runs completed may be found in 

Appendix B. 

As might be expected from the random nature of the ship motion, 

the time_ to complete the task was the most variable parameter. The 

average' time over each series of 5 runs as a function of sea state, 

HUD, and T/W (Figure 14) lies in a band from 19 to 36 seconds. 

0 

0 

e:. 

0 

0 

TEST 

CONOlTION 

SSO IIUD3 V.C. 

Ss4 IIU03 V.C. 

SSG IIU03 V.C. 

SSG PoUOI V.C. 

ss6 HUD3 A.C. 

}IEAN 

FLICHT 

TIm: 

(sec:) 

50.0r-----------------~-----------------r~ 

40.0 

<0 0 8 

30.0 0 C:> 0 

~ 0 0 

0 

9 0 6 

20.0 0 0 

10.0~----------------~--~--------------~~ 

1.00 1.0:; 

Tl\RUST-TO-WEl~IT RATIO, T/W 

1.10 

Figure 14: Mean Flight Times for the Piloted Simulation 

Individual times varied from a low of 7.2 seconds for a particular 

sea state 4, HUD3 run, to a hlqh of 79.0 seconds for a sea state 6, 

HUD1 run. It should be noted that the addition of the sidetask does 

not appear to change the amount of time the pilot takes to complete 

the task. 

38 

. 

-.. 

! . 

I~ I , 

... 

-~ , 

f .. I 

- \ 

I 

j 

i 

PI-
The non-piloted simulation showed that the mean time and 

standard deviation for flight time did not approach a steady value 

until after approximately 40 runs, indicating why such a large 

variation exit~ for the piloted simulation gro~ps of 5 runs each. 

Average touchdown velocity as a function of T/W for HUD1, sea 

state 6 conditions are presented in Figure 15 for two values of 

engine response. The following points should be noted: 1) In- , )0.,. }. 

general there is a l~ss than expected change in touchdown" velocity" i 

I 

with change in T/W ratio (less than 4.0 ft/sec for the T/W ratio I 

range tested), 2) The faster engine time response (0.3 sec-s) 

produce touchdown V6locities of 0.5 to 1.5 ft/sec less than the 

slower responding engine (0.7 sec;s) and 3) The greater vertical 

velocity damping the lower the touchdown velocity (as much as 4 

ft/sec) • 

pilot ratings as a function of T/W for IruDl, sea state 6 

conditions are presented in Figure 16 for two values of engine 

response. The following items should be noted": 1) The pilot 

rating is higher (worse) as T/W ratio goes down over the range 

tested, 2) pilot r~ting is higher (worse) as engine lag is 

increased and as vertical velocity damping is decreased (as much as 

2.5 points difference) 3) l'lith the IruDl system and low vertical 

damping there are ratings in the Inadequate range even for T/N 

ratios of 1.1. 

Average touchdown velocity and pilot ratings as a function of 

the control system and HUD format are presented in Figure 17. 

39 

\. 

'-
I 

I 

(. 

Z" 

s{'c-1 

0 0.0 

A -0.2 

0 -0.4 f: 

~,ibA_'·".m 

20.0r---------~----~~--------------~-, 

1.5.0 

HUDI 

Sea St2'te 6 

Velocity Comd. Sys. 

T eng - 0.3 sec 

0 

Q 

0 

T 

TOUClI-oo;m 

VELOCITY. ~T) 

(ft/sec) 

10.0 

5.0 

0.0 1.--'-______ --,. ____ .1. ________ ..._-------'--' 

1.0 1.05 

Z" 

sec-1 

20.0 

0.0 r -0.2 

HEA.'1 

-0.4 

}lIN 

15.0 

eng - 0.7 . sec 

10.0 

TOUClI-OQ;m 

VELOCITY, ~O 

Cft/sec) 5.0 

0.0 

1.0 

(. -> 

TIlRVST-To-WELcnl' RATIO. TN 

HUDI 

Sea State (, 

Velocity Comd. 

( ! (;> 

1.05 

T1IRUST-TO~~EICRT RATIO, Tty 

1.10 

T 

Sys. 

~ . 

~ 

1.10 

Figure 15: Influence of Thrust-to-\'/'eight Ratio, Engine and 

Airframe Dynamics on TOuchdmm Sink Rate--

Baseline HOD with Velocity Command Control System 

40 

-

f __ -

, 

'.-
(', 

U 

I 

r--

r 

o !l.0 

6 -0.2 

o -0.4 

f~ 

.3. = sec 

In adequate 

PILOT RA TING 

Ad 

Sa 

o 0.0 

"\J -0.2 

o _-0.4 

I 

equate 

t1sfa~tory 

f:: 

.7 sec: 

n:.deQuate 

PILOT RA TING 

Ad equate 

10 

9 

8 

7 

6 

5 

4 fJ 

2 

1 

1.0 

10 

9 

8 

7 f6 

5 

4 

J 

s atisfactoty 

2 r 

1 

1.0 

(. 

-

-

t 

-

HUP1 

Sea State 6 

/ 

/ 

Velocity Cot>d. Sys. 

0 

-r - - -~- -

~ 

- t_ - - - - -

LOS 

THRUST-TO-WEIGHT RATIO, T/~ 

HUPI 

Sea State 6 

.. " 

0 

- -

.;. 

c" 

G 

- -. 

I 

1.10 

<) :Velocity COl:!d. S)ls. 

-

-

f -J- -, -

- - - - - -

LOS 

ntRUST-TO-WEIGlIT RATIO, T/'rl 

<) 

-"\J 

0 

- -

1.10 

Figure 16: Influence of Thrust-to-weight Ratio, Engine 

and Airframe Dynamics on pilot Ratings--

Baseline HUD with Velocity Command Control System 

41 

.:;. 

-, 

f' 

--

2' 

! 

::-..: 

~~[ '-. 

~:: .I 
.~, 

0 

c. 

0 

.. /0/.-' 

/ . 1 I 

I / ,,/ , . . .' 

Sys.1. 20.0r---------------~----------------~_, 

HUD 

V.C./1 r V.C./3 MEAN 

A.C./3 15.0 

~N 

TOUCH-DOWN 

VELOCITY, ~D 

(ft/sec) 

10.0 

5.0 

0.0 

10 

1.0 

Sea State 6 

Z - -0.2 sec-1 

T

W - 0.3 sec , eng 

I 

1.05 

11lRUST-TO-\lEIGHT RATIO, T/W 

1.10 

PILOT RATING 

Inadequate 

Adequate 

Satisfactory 

9 

8 I7 

6 

5 

4 

31-

2 

I -- -

o 

o 

flL-----------~-----~·----------------~~ 1.0 1.05 1.10 

THRUST-TO-'-IEIGlIT RATIO, TN 

Figure 17. Comparison of Touchdown Sink Rate and Pilot 

Rating as Influenced by Thrust-to-Weight P~tio 

for a Selected Condition--Baseline lruo with 

Velocity and Attitude Control Systems 

42 

.' 

,',, , 

x. 

i ';!::'~ 

j T)'~ . '. ~ 

, 

" 

I 

I 
.~. " 

. , 

) 

.' I 

: ,. / . I ,. 

\./ 

As shown, the touchdown velocity and pilot ratings are both improved. 

with the additional information of the HOD3 format. There is as 

much as a 2.5 ft/sec improvement in average touchdown velocity in 

going from HUDl to HUD3. It should also be noted that HOO3 produced 

an Adequate pilot rating throughout the range of T/W ratio tested 

for the stated engine response and vertical damping (Figure 17). 

The addition of the side task did not seem to influence the result in 

any systematic manner. 

Average touchdown velocity as a function of T/W for Iruo3, sea 

state 6 conditions are presented in Figure 18. Again note that 

there is li ttle change in average touchdown ve.1oci ty over the T/W 

ratio range tested (less than 3 ft/sec for the worse case). Average 

touchdown velocities are 0.5 to 3 ft/sec better with the Hun3 format 

than for the HOOl format (Figures 15 and 18). The engine time 

constant and vertical velocity time constant have a much smaller and 

less consistent effeGt on average touchdown rate for the HUD3 format 

(Figures 15 and 18). 

pilot rating as a function of T/W for 1ruD3, sea state 6 

conditions are presented in Figure 19. The most interesting thing 

to note here is that the average ratings are Adequate for all 

conditions tested throughout the T/W ratio range tested. The pilot 

ratings also. show more clearly that the faster responding engine and 

greater vertical velocity damping are important to the pilot as 

evidenced by the consistent effect they have on pilot rating. There 

is as much as one pilot rating improvem~n~ attributable to either a 

fast responding engina or good damping. 

43 

J' 

, 

./' 
I 

. .,. /' !, I 

/,; j.. 

I ' I . /, 

, / 

1 \' /1' 

',/: \--

0 

0. 

0 

T 

0 

'i] 

0 

T 

z 

w 

"",,-I 

0.0 r -0.2 

-0.4 ~ 

MIN 

eng - 0.3 sec 

TOUC!t-DOYN 

VELOCITY, ~D 

(ft/sec) 

z v 

sec-I 

0.0 r -0.2 

-0.4 MU~ 

HI~ 

eng - 0.7 sec 

TOUCH-DOWN 

VELOCITY, ~ 

(ft/sec) 

~. I , . / 

20.0r---------------~~----------~--~_, 

HUD3 

Sea State 6 

15.0 Velocity Command Sys. 

10.0 

5.0 

0.0L---------------~~--------------~~ 

1.0 1.05 

THRUST-TO-WlrIGH:r RATIO, T/~l 

1.10 

20.0r-------~--------~--~--~----~----~~ 

HUD3 

Sea State 6 

15.0 Ve1o€ity Co~and Sys. 

10.0 I-

.> 

() (, 

5.0 

O.OL---------______ ~~ ______ --------~-J 

1.0 1.05 

THRUST-TO-WEIGHT RATIO, T/W 

1.10 

Figure 18: Influence of Thrust-to-weight Ratio, Engine and 

Airfra~e Dynamics, on Touchdown Sink Rate--

Augmented HUD with Velocity Command Control System 

44 

'! -:,-' 
.l- I .. 

;,." . 

" 

Z 

II 

sec-1 

D 0.0 

Do -0.2 

0 -0.4 

T 

en - 0 

In 

?~ .. 'lEA!' , 

:11~ 

.) sec 

adequate 

PILOT R.\ TING 

equate 

\" 

S atisfactory 

o 0.0 

'i) -0.2 

o -0.4 

Ten.g ., 0 

?~X HE.A.'l 

Ml~ 

.7 sec 

In adequate 

PILOT RA TING 

equate 

Sa tisfactory 

10 ...:... 

9 

8 

7 

6 

5 -

4 

3 

2 

1 

1.0 

10 

9 

8 

7 

6 

5 

4 

3~ 

2 

1 

1.0 

. HUD) 

Sea State 6 

Velocity Co~and System 

t -{ - - - - -

% _1 - - - - - -

1.05 

TIlRUST-TO-viEIGHT RATIO, T/W 

HUD) 

Sea State 6 

Velocity Co~~and Systeo 

t -t -f - - -

\ - - - - - - - -

L05 

T"rlRUST-TO-I-1EIG!!T RATIO, T/W 

I 

. - -

-, 

I 

1.10 

- -

() 

-~ 

1.10 

Figure 19: Influ~nce of Thrust-to-weight Ratio, Engine and 

Airframe Dynamics on pilot Ratings--Augmented HUD 

with Velocity Control System 

45 

'. 

. :. t~.·.'~.~~· (: . 

..--

. -.~-~;;. 

'\. ; 

~,,'::~ 
" /' , i ,.-----' -/ 

-! II 

-r" ,-

_.t'- ' 

Figure 20 provides some perspective on the effect of sea state 

on average touchdown veldcity and pilot rating, for a sel~cted, 

condition using the HUD3- format. Note ,the relatively small change ~>~- -----L I "~ i I 

in touchdown velocity over the range of sea state, the greatest 

change occ~rring for T/W ratio of 1.01 (an increase of 5.0 ft/sec I 

from calm sea to sea state 6). The pilot ratings show satisfactory 

ratings for sea state 0 8nd 4 and Adequate ratings at sea state 6 

for all T/W ratios for the selected condition. 

The miss distance when performing the sidetask associated with 

the attitude command system is presented in Figure 21 for various 

values of T/W. The miss distance is the horizontal distance from 

the aircraft center of gravity to the center of the bullis-eye on 

the ship deck at tHe time of landing. It is interesting to note 

that the average miss distance changes less than a foot with the 

fast responding engine while changing 5.5 feet for the slow 

responding engine over the T/W ratio range tested. In both cases 

the miss distance decreases with decreasing T/W ratio. Possible 

reasons for th'is result would include; too much control sensitivity 

at higher T/W, or a change in piloting technique; i.e., the pilot 

may be spending a different percentage of time, or changing the .... 

frequency with whi~h he samples the information at the different T/W 

ratios. ,-/ 

A better understanding of the relative performance of HUD1 and -~/- I 

! HUD3 , can be gained from Figure 22, which shows a histogram of the 

landing performance for HUD1 and HUD3 with a histogram of the ship 

deck motion for sea state 6 cond~t~ons. 

46 
/ ./ 

/' . . ' .,~ 

.. ------

!: f/" ' .. ,t ./ 

I' 

I 

, ' I .,; " . , ',/"Y i,.'" , '. ;)'>" ,'" '" / .... ; ,'. '. . I ' ./ .. • ",,;. I ,,~,:. • • 

0 

6 

0 

TN 

l.01 r l.05 , ~lEA.'I 1.10 

MIN 

TOUCH-DOlIN 

VELOCITY. hTIl 

2G.0 

15.0 

10.0 

(ft! sec) 5.0 

0.0 

10 

PILOT RATING 

9 

Inadequate 8 

7 

6 

Adequate 

5 

4 I3 

Satisfactory 

2 

1 

/' / . V 1'" 

I I" 

/1 I, \ ... :1 'I 

HUD3 

VELOCITY COI1MA.'1D SYSTEM 

Z • -0.2 sec:-1 

TV 

eng • 0.3 sec: 

~ CALM 2 

SEA STATE 

'1 

.. ' 

--- . 

t 

4 

,. " • / -! 

I. ,. I • ~ . I ' oJ', ; , 

(, 

6 

----'-----

------(3-

(3 

o 

CAUl 

o 

2 4 

SEA ST,ATE 

6 

Figure 20: Co~parison of Touchdo\ffl sink Rate and Pilot Rating 

as Influenced by Sea State for a Selected Test 

Condition--Augmented HUD with Velocity Command 

Control System 

47 

, / ./' 

1 :,' 

I~ 

Itt;. i 

i t; 

.--r-' 

; .~ .~ 

I," 

';: 

',.,:. . '..,;. 

i ';~~ 

'J 

/ 

/ 

/ 

I '. > 

. 

I 

I 

._---!-:-: 
TEST 

CONDITIONS 

0 'teng.- 0.3 

6 't 

eng - 0.7 

SS6 IIUD3 A.C. 

HISS 

DISTANCE 

eft) 

Figure 21: 

IIISTOGRAMS 

-- Landings 

using RUDt 

-- Landings 

using RUD3 

..•. Ship deck 

cot ion 

PERCENT 

OF 

TIME 

;. 

·1 

16.0r--------~-------.....-. 

14.0 

12.0 ~ 0 

0 0 

10.0 A 

8.0~--------------L---------------~ 1.0 1.05 

TIlRUST-TO-WEIGRT RATIO, T/W 

1.10 

Influence of Thrust-to-weight Ratio on the 

Positioning Side Task--Augmented ~~ with 

Attitude Command system 

3.0 

2.0 

1.0 

I 

I 

I - ............. I 

I 

, 

/ 

r·o 

I 

I 

/' 

I \ / \ 

I \ / 

0.0~-4~~--~--~--~~~~w-~~~--~~ 

0.0 10.0 20.0 

TOU01-D~riN VELOCITY AND SHIP DECK VELOCITY (ft/sec) 

Figure 22: Distribution of Touchdown Sink Rate and Ship 

Deck velocity--Baseline and Au~ented HUD 

Breakdown for all Landings Completed for 'the 

Simula tion 

48 

~ '. , . 

-!' ! 

. ;; .. ;/ .' J ~.' 

:' 

'" / , , ;' 

I' 

, 1 .• 

.' .-~; , . 
The velocity histogram for HOD 1 peaks at 8.5 ft/sec, which is 

just above the 1 sigma value for the ship motion velocity for sea 

state 6. The tc)uchdown velocity for HOD3 peaks at 6.5 ft/sec or 2 "' 

ft/sec less than HOOt. The histogram also shows the HOD3 data to 

have a smaller standard deviation. 

Table 3 indicates the precentage and number of landings over 12 

ft/sec which occurred overall and for several sets of test 

conditions. 

Table 3: TOuchdown velocity statistics Indicating Number 

of Landings with Sink Rates Greater Th'qn 12 ft/s 

for Selected"~st Conditions. 

Nt'MllER OF 1. 

CONDlTIOlI ~mrnER OF LANDlHCS LA!mI~CS 

LA:mINr.s > 12 ft/sec:. > 12 ft/sec:. 

Total S"il:mlation 995 28 2.81 

5.5.0, 5.5.4 Overall Z60 0 0.0 

5.5.6 Overall 735 28 3.81 

5.5.6, RODI Vel. C:nd. Sys. 212 18 8.49 

5.5.6, HUD) Vel. Cmd. 5ys. 419 7 1.67 

5.5.6, HUD) Atd. Cmd. Sys. 104 3 2.89 

5.5.6, HUD3 V.C. T 

eng 

-0.3 216 1 0.46 

5.5.6, HL'D3 A.C. T eng 

-0.3 60 1 1.67 

"" 

11. series of T/W ratio histQgrams are presented in Figures 23 through 

27. Most of these curves show a fairly sharp spik~d peak. In 

observing the simulation and through pilot comments, it is" concluded 

49 

:. 

. 

I 

! 

1 

i 

i 

I' 

I 

...J." 

I" 

") 
/. " ! f; . 

. , .... / 

(', 

that this shape of curve can be attributed to the ~ilots technique 

of setting up an initial descent rate and holding it either un·til 

near the ship deck for the HUD1 system, or until the 3 sigma line on 

the HUD3 system was approached, and then slowing or arresting the 

descent. The initial descent phase, which takes 90 to 95% of the 

total time produces the predominant peak in the histogram. 

The histograms for HUD1, sea state 6 conditions are presented 

in Figure 23. 

40;0 

30.0 

PERCEllT 

OF 

TIME 

20.0 

10.0 

Figure 2: 

RUnl 

~ea State 6 

Velocity Ccd. Sys 

Z - -0.4 sec-1 

TV • 0.3 sec eng 

T/Ilmax 

01.10 

01.05 

61.01 

TIlRUST-T04::1GHT RATIO, TN 

Influence of Thrust-to-weight Ratio on 

Thrust-to-Weight Ratio Use Histograms--

Baseline HUD with Velocity Command Control 

System 

The peak occurs at the T/H ratio used most often in the initiai 

descent. The peak is somewhat broader for T/Wmax of 1.1 probably 

50 

; , \. I 

I /' . 

/ /. 

,,' 

I 

~ , 

1 

I 

j 

I 

,. 

, . 

i 

i 

! 

/ 

(J; 

" /' , 

I i 
I' 

- , __ .1' • ... ;'-, . 

... , M M .• 4. ,:. 8I!ii<'JCl • II •• 

indicating that the pilot is operating with lower gains. The base 

of the peak is fairly narrow and correspond~ to the observed 

pilot behavior, with the HOD 1 format, of chasing the deck less, 

possibly due to lack of positional cues. Landings, using the HOD1 

format generally occurred as the deck caught the aircraft. The area 

under the curves tends to bunch up at the high end of the available 

T!W as T!Wmax is decreased. ~nis is due to the pilot tending to fly 

more conservatively (lower descent velocities). 

The histograms for HOD3, sea state 6 conditons are shown in 

Figure 24. 

SO.Or---~--~----~--'---~----r---~--~----r---~ 

40.0 

Ruro 

Sea State 6 

Velocity Ccrmd. Sys. 

Z - -0.4 sec-1 

TW • (\.3 sec eng 

T/'rlmax 

01.10 

01.05 

Co 1.01 

30.0 

PERCElIT 

OF 

TIME 

20.0 

10.0 

o.o~~=8~~~L-~--~~~~:O==~=C~ 

0.90 0.98 1.04 1.10 

THRUST-TO-WEICHT RATIO, T/W 

Figure 24: Ir.fluence of 'lbrust-to-vleight Ratio on 

Thrust-to-Weight Ratio Use Histograms--

Augmented HUD with Velocity com~and Control 

System 

51 

," 

, ~.' 

· ."; 

· . 

· , 

, .: 

1 .~ · .-

,-

.. <. 

'\, " t~ ,..-. . it 

-* :~'i .:: .. 

./ 
I 

! 

Notice that the curves broaden somewhat in comparison to those 

for the HUDr format. Because of the positional informa·tion 

available to the pilot with the HOD3 format, pilot gains are 

somewhat lower, and the pilot has more opportunity to chase or run 

from the ship deck. This tends to broaden the peak and base of the· 

curves. The area under the curve for T/Wmax = 1.01 is much more 

bunched up at the high end, than for the HUD1 case. This is due to, 

the pilot's tendency to fly more conservatively (lower descent 

veloci ties, higher thrust settings), si nce the· HUe3 ~ormat makes him 

more aware of the limitations of the available control power. 

The effects of vertical velocity and engine time lag for HUe3, 

sea state 6 conditions ar.e presented in Figure 25. 

With zero vertical damping, the pilot had a much more difficult 

time controlling vertical velocity, and this is compounded with a 

slow responding engine. The curve becomes much les~ peaked and is 

spread out over a larger range of available T/W. Thi~ is in 

contrast to the curve for higher damping and faster responding 

engine where the pilot is able to control vertical velocity with 

much less th.rottle movement (i.e., the throttle becomes a vertical 

velocity command control). 

The effects of HUD forma t and type of control system on T/W 

histograms is presented· in Figure 26. M interesting thing to note 

here is that the peak for the attitude command system occurs at a 

higher T/W ratio than for the velocity command sY5tem u!'ling either 

HOD. Also the curves for thE; attitude command system tend to be 

52 

.,~ -

, 

I 

I 

I 

I 1 

1 . il 

I ~ 1 

I 

! 

I 

i i. 

'i 

'I~ , - ~ .; 

i " 

.... :.~ .} . .-
I 

! _.J 

sO.Or---,----r-~~--~--_r--~----T_--~--~--~ 

. 40.0 

PEll-CENT 

OF 

TIME 

]0.0 

20.0 . 

10.0 

RUD] 

Sea State 6 

Veloctty Comd. 5Y8. 

~. 

f \ 

I \ 

. 1 

Z 1: TIs'" w eng max 

sec-1 sec 

0-0.4 Q.3 J::l 

o 0.00.7 1.1 

A 0.0 0.7 1.01 

O.o'-i~~~:::"'~_L-_lJ::c~~bt:;;E(b:=:fi:d 0.90 0.98 

. ~.~.-: i.04 1.10 

TIIRUST-TO-UEICHT RATIO. TIll 

Figure 25: Influence of Engine and Airframe Dynamics, 

40.0 

PERCENT 

OF 

TIME 

30.0 

20.0 

10.0 

and Thrust-to-Weight Ratio on Thrust-to-Weight 

Use--Augmented HUn with Velocity Command Control 

System 

Sea State 6 

TIll • 1.1 max Z • -0.4 sec-1 

TV • 0.3 sec eng 

RUn 

01 

o 3 

A 3 

SYSTE:-I 

Vel. Cmd. 

Vel. Cmd. 

Att. Cmd. 

TIIRUST-TO-UEIGIIT RATtO. T/W 

Figure 26: Influence of HUD Format and Control System on 

Thrust-to-Weight Use for a Selected Test condition 

53 
GI'$;_lHfr~~"h ' ~ .:_:::: ,I _~.- .~~ .......... --., .:/ / .r .. : ,-~ " ,/ , --.... ---. ' 

r~ 

much more symmetrical abou~ the peak. The explaination for this is 

that the pilot was observed to spend at least as much time on the 

positioning sidetask as with the vertical task. As a result, he 

tended to fly mor9 conservatively (at a higher thrust, slower 

descent rate) and th~ errors in desired vertical position/velocity 

are therefore more likely to occur randomly rather than only on a 

conservative side as when flying only a vertical task. 

The effect of sea state on T/W histogralllS for a selected 

condition is presented in Figure 27. 

As expected, at the low sea states the pilot spends more time 

in the initial descent and less ti~8 correcting for deck position. 

As a result, lower sea states produce a histogram with a sharper 

peak and with a '~maller base. 

SO.Or---,----r--~----~--~--,_--_.--~~--~--, 

40.u 

P£R~ "J 

OF 

Tl~ 

20.0 

10.0 

HUD) 

Veloc1t"y Co!!!d. Sys. 

T/U • 1.1 max 

Z • -0,4 sec-1 

TV • 0,3 sec eng 

Sea State 

A 4 

o 6 

o. 0 l=l~Q~~~L.-_1.---iL...-.-J~:;&!=:C~:::::f~~~ 

0.90 0.98 1.10 

'I1lRUST-TO-:;ZIGllT RATIO, TN. 

Figure 27: Influence of Sea State on Thrust-to-Hcight Ratio Use--

Augmenteu HUD with Velocity Command System 

54 

" 

~~~ --- .. ,t~ . 

'-:',' 

~. 
tn-Ill- - +";e'''. e • 

eo 

.~; . 

2. pilot Comme~ts 

The pilots were given an explanation of the task and 

discription of the HUD s~bology. From this they developed some 

individual techniques. Pilot A, especially for the higher values of 

T/W· with fast responding engine, often waited for the deck symbol to 

crest just below the 2 sigma line., and then smoothly rolled off the 

throttle providing a quick descent in which the aircraft would catch 

the deck on its downward motion with a usually low value of relative 

descent rate. pilot B used this technique also although, not as 

often. 

Both pilots became more adept at picking out "lulls" in the 

ship motion as the simulation prqgressed. It was often possible to 

tell that the deck was in a lu~l condition when its motion was slow 

and position was a couple of feet above the mean line. The pilots 

would make a quick descent and attempt to catch the deck before the 

more extreme motions reoccurred. 

The pilots also understood that if they chased the deck below 

the mean line (or were unable to arrest a descent until below the 

mean line), then in most cases it was better to continue and attempt 

to catch the deck n·ear its lowest position (and therefore low 

veloci ty) than attempt to climb back to the 2 sigma line. \-1hen the 

attempt to pull up was made, especially with low T/H, the deck 

tended to catch the aircraft near the mean deck position, often with 

a high velocity, and therefore, high relative velocity. 

55 

-.. '~ ... ~ ··vit1¥Z&1!P 

. ' 

" 

'1M'. 

."" '" .. 

.. ~ 
(. 

I / 

I;;.......;..;;~_, _________ .;.;;··~· ~·."'_."'''~ ... '.......... m ... __ .. ~ 

." 

'N' t"C53C?_r' r "3T7aam--'-- .... -....... ...... - , ........ ,~-

/ 

.' 

-"-l ,".-.. 

M ------..~<!.. 

----- --~ ... 

I 

/ ..... " .. 

The most general comment expressed about technique was, "always 

be in a position to gradually take off power; and don't get caught 

needing it." 

Pilot comments indicated the following: 

1) The greater the T/W, the more controllable and easier it is 

to perform t.'le task (through the range tested). 

2) The higher the value of T/W, the less sensitive the pilot 

workload is to engine lag and vertical damping. 

3) The higher values of damping provided better control of 

vertical velocity, which in turn aided the initial descent to the 2 

sigma line. The lower values of damping provide q'.licker vertical 

response and therefore greater agility during the final phase of 

descent. 

4) The slow responding engine was considered unfavorable, even 

though a few cases occurred when the engine compensated for an 

over~ontrol by the pilot. 

5) Both pilots commented on the importance of engine noise as 

a cue. 

6) HOD3 was perferred over Iron1. Th~ pilots commented 

favorably on having the ship motion boundaries and 2 sigma lines as 

references in giving precise situation information. 

7) Comments concerning HUD' generally focused on the feeling 

of not knowing either the position or vertical velocity of the 

aircraft relative to the mean deck position. 

56 

.,.,-:-.... . ".-

. i 
I 

;" 

,. 

B) When using HUn1, the effects of damping and engine lag were 

less evident, since errors were not as easily detected due to the 

lack of references. One of the pilots commented that the task 

workload was less using HUn1 because of the lack of references, but 

gave it a higher (worse) pilot rating because of the greater 

uncertainty involved. 

57 

; . 

-~. 

! 

\ 

I 

I 

I 

, 

r 

\ 

i-

~ ~ . ,.-

; .• 
..... 

IV. REVISED ANALYTICAL PREDICTIONS 

After the piloted simulation was run, the non-piloted 

simulation was modified in an attempt to match the measured T/W 

histograms, as well as to better represent the observed piloting 

technique. 

A. MODIFIED ANALYTICAL MODEL 

The following modifications were made to the program: 

1 ) A feed back loop was added to provide vertical veloci ty 

damping through thrust, and the coefficient for vertical velocity 

damping through airframe was reduced to more closely match the 

fixed-based simulation model. 

2) The flight path command logic was rewritten to provide a 

better match to the observed landing strategy adopted by the 

pilots. Details of the logic are given in the next section. 

3) Two additional noise sources were added to account for 

pilot perception er=or an~ internal pilot noise. 

4) The pure ti~e lag, used to represent the pilot's 

information processing time interval, was divided int~ a pure lag 

and a seconda~y time in which the input to the pilot's neuromuscular 

dynamics was held constant for a specified time. The purpose her~ 

was to simulate the pilots concentration on a side task. A block 

diagram of the modified model set-up is shown in Figure 28. 

58 

l~ I {'<;" 

. ~: 

"~~~ fJ 

I 

• i~~.i 1"1"1- I ~';'; 

If'} \ II:"' 

I !F~:. , :r.t~ I 11'0 

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: \ (">.:f- • _____ _ 

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• .;J ,~>~ 

~ '" 'C/, 

It(~. -
') . 

V1 

\0 

./ ,,' '- , / 

: ."1~ ," /' 

r--

I 

L 

PILOT HODEL 

FLIGIIT S 

PATH 

+ LOGIC 

SHIP 

tlOTlON 

tlOor:L 

PILOT 

DYNAMICS 

Kp (TL II + 1) 

(Ie s + 1) 

SEA STATE CONDITIONS 

-, 

I 

--' 

) 

TURBULENCE 

HODEL 

SEA STATE CONDITIONS 

Figure 28: Block Diagram for the Modified Non-Piloted Simulation Model 

\ \ \ 

\ \. 

\ 

\ \ 

-) 

\! 

I,. '\ -. --. 

\ 

., ·t 

" \ 

, \ \ '. 

\ 

,\ "". , 

" 
I 

~ 

i 

fO 

" ! i , ;/ I / /.. / , r l/ /" 

.. ' ,.~ i- i .. .' 

" ! I 

;.. i :-- I / , ! .. .:'~ i . .~: i I ./ of· 

The modified version of the flight path command logic consists 

of 6 basic sections. The first section determines which of five 

defined regions of relative position and velocity that the aircraft 

is currently in. This section then specifies which of the remaining 

5 sections is to be useq to supply the commanded flight path 

information to the pilot transfer function. These sections are 

referred to according to their basic strategy: RUN FAST, RUN, 

CHASE, ABORT TO HOVER HEIGHT, and CHOP THROTTLE. The RUN FAST 

sequence is initiated whenever the relative velocity of the ai~craft 

and ship exceeds a given value, typically 5.5 ft/sec. The RUN FAST 

logic commands an altitude of 12 ft above the present altitude using 

a cosine smoothing function. This command causes the simulated 

pilot to apply full throttle. If the relative position is l.ess than 

9 ft and the relative velocity eKceeds 4.5 ft/sec the RUN sequence 

is initiated. The RUN logic commands an altitude of 6 ft above the 

present altitude through a cosine smoothing function, causing the 

simulated pilot to apply approximately 95% of full throttle. If the 

aircraft was' in the RUN FAST sequence. and then switches to the RUN 

sequence, the commanded altitude is lowered 6 ft from the previously 

commanded altitude, thus causing the throttle to be reduced from 

full to approximately 90%. If the relative position of the aircraft 

and ship deck is less than 3 it and the re la ti 'Ie ve loci ty hasn't 

exceeded 4;5 ft/sec then the CHASE sequence is initiated. The 

CHASE logic commands a cosine function descent, modulated by the 

ship deck motion, starting at the previously commanded altitude and 

60 I', 

I '. 

!:~< 

, 

"-Ii 

. . 

.~ 

/ 

! -' 
/' 

'-" -t 

? , 

( 

--\ 

.. ---- ..' " 'I I," ,;. 

.. ' '," 

,-

/' ~,' 

~.~ .< '/ ," I / 

ending at the ship deck. If the aircraft position descends below a 

designated abort height, the ABORT TO HOVER HEIGHT sequence is 

initiated. The abort height is the height above the ship deck mean 

position which, if a descent is continued, will not provide enough 

time to gain the necessary height to prevent a hard landing in the 

event the next segment of the motion is around the 3 si~ value. 

This height, of course, varies with sea state. Up to sea state 4 

the abort height is zero because the relative velocity can be 

maintained below the gear limits anywhere in the ship motion 

boundaries. The hover height is the 2 sigma height above the ship 

deck mean. The ABORT TO HOVER HEIGHT logic commands the hover 

height altitude through a cosine smoothing function from the 

altitude in the previous sequence. The hover height altitude is 

then maintained until conditions require use of another logic 

section. The CHOP THROTTL~ sequence is designed to mimick a normal 

pilot landing technique. It is initiated when the relative aircraft 

to ship position- is less than t.~e Chop Throttle Now Height (CTNH), 

which is another adjustable variable. Unlike the other sequences, 

which can be abandoned for a more apporpriate one at any time during 

the sequence, once the CHOP THROTTLE sequence is initiated it 

continues until a landing occurs or a specified time has elasped 

The CHOP THROTTLE logic commands a 12 ft descent in altitude through 

a cosine smoothing function. This command effectively produces near 

zero thrust output. The sequence maintains this reduced altitude 

for a specified period of tim~. If during this time period a 

61 

, 

i 

I 

; 

I 

, 

I 

I 
, 

l 

1 ,: 

. .' . 

. r--.. • 

./ 

/' ., 

______ ....... _.....;.·'::.:_~:....l- ... ' " ',' 

/' 

./ .. ",.j' 

landing has not occurred, the sequence commands an ascent back to 

the hover altitude. This logic models the pilot behavior after he 

sees a landing opportunity; i.e., the aircraft is 0.5 ft off the 

deck but with the deck beginning to descend. 

The modified analytical model produced a much more accurate 

representation ox observed piloting technique. For example, it was 

observed in the piloted simulation that. with the slo.w responding 

engine, more landings were made running from the ship than chasing 

it, whereas with the fast responding engine, approximately the same 

number of landings were made ch~sing·the ship as running from it. A 

series of computer runs were made which duplicated this result 

(Table 4). 

Table 4: Influence of Thrust-to-weight Ratio and Engine and Airframe Dynamics on Flight Path Command Logic Sequence Use 

TEST COllDlTION ~T OF LANDWCS 

T/'rl T Z Flight Path logiC Section 

eng w 1 2 6 7 sec sec-1 

1.07 0.7 0.4 52.5 10.0 35.0 2.S 

1.05 0.7 0.4 55.0 10.0 15.0 20.0 

1.03 0.7 0.4 55.0 5.0 27.5 12.5 

1.01 0.7 0.4 50.0 15.0 30.0 5.0 

Average 63.1 36.9 

1.01 0.3 0.4 42.5 10.0 27.5 20.0 

1.07 0.3 0.4 42.5 5.0 32.5 20.0" 

1.05 0.3 0.4 40.0 12.5 30.0 17 .5 

1.03 0.3 0.4 45.0 2.5 45.0 7.5 

Average 50.0 50.0 

Flight Path Logic Sections: 

1 - Run Fast Sequence 

2 - Run Sequence 

6 - First section of the Chop Throttle Sequence 

7 - Second section of the Chop Throttle Sequence 

62 

.I !.' 

f· . 

f' 

I ! ,. 

I 

.. , . 

! 

/ 

,. ! 

I 
/ 

." ,':- ! / / .• to 

" -.- ~ -.-

r,~ , ~ I 

I "; The original flight path command logic did not provide a chop "! 

.. -,~ 

throttle sequence, and ";; the landing nearly alw().ys occurred while in 

! ,~ 

< 

~~~ the RUN FROM sequence. A schematic diagram of the modified flight 

! ~: '. .~ path command logic ~s shown in Figure 29. An example time history 

showing some of the flight path command logic aspects is shown in 

Figure 30. In addition,. two representive time histories are 

presented in Figures 31 and 32. 

B. CQMPARISON OF ANALYTICAL AND· SIMULATION RESULrS 

A comparison of touchdown velocities for the non-piloted 

simulation and the piloted simulation are shown in Figure 33. The 

non-piloted simulation results, using either the original or 

modified flight path command logic, produced lower touchdown 

velocities than were achieved ia the piloted simulation. The 

modified logic ge'nerally produced the lowest touchdown velocities. 

Further adjustment of the parameters in the non-piloted simulation 

to produce a closer match to the piloted simulation results could 

provide further insight into the pilot's capabilities. 

The results for the modified version of the flight path command 

logic when compared to the piloted simulation data show a good 

correspondance in average flight time for the lower T/W ratios, but 

./ 

a large gap is evident for T/W = 1.1 as shown in Figure 34. 

63 
C\ 

"" 

) 

YESI 

IS 

IIAW IFLAG1 

PROGIWI ~ • 1.0 

AND TE4 

< 1.5 

~ 

~ RUN FAST 

SEQUENCl; 

INITIALIZATION OF 

LOCIC fLAC;S: 

GTrI1l: TIrlE COl/-

STANT, II/ITIAL 

POINT, AND AHPLlTUDCS FOR COSINE StlOOTIIING 

FUNCTIO:I. 

IS 

~ NO SS2 - ZD 

> 4 5.S. 

I YES 

.. ~ RUN 

SEQUENCE 

" 

;:::1 PI r 

IS IS 

SSl - Z NO SSl- Z 

> - 9.0 4 < -3 

AI/D 

SS2 - ZD 

> 4.5 I YES T YES 

! 

ABORT TO HOVER 

HEIGHT SEQUENCE 

" 

, 

----. .. ,. 

NO 

.--, ." . ~ ......... ~ 

IS 

Z 

4- < ABRTII 

T YES 

J 

CHASE 

SEQUENCE 

" 

IS 

NO SSl - Z 110 

4 >CTNH -- I YES 

! 

CHOP TIIROTn.t 

SEQUENCE 

" 

'-- - -- ---- -

- --- ----

figure 29: Block Diagram of the Modified Flight Path Command Logic 

'''\':2lr:\_::tf~ ol~ ~~5;,'i<:i:.~:§~:;J~L;:i';~~r.;:U;t~~·; \}", !,~ t~\ J"';~' .,( i ~ ',' ·,,,(,,·::~,·,~~:.~·_~·~:\·~~y3: ~ ~.tr,: ·:~~~.~_.f!·;~'.:· __ ~.~,~'_~1;·:?:·' ,.,;'-_~~~~ 'i,~,~.: .:};;,";;;.;,: .. ~~;;J:~:';i,:v \ 

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59 ' 

INITIAL LET DOWN TO HOVER 

,HEICHT 

lIAIIlTAININC HOVER IlEIGHT 

Relative velocity > 5.5 

ft/sec , R~~ FAST SEqUENCE 

A/C < 3.0 ft above deck 

CHASE SEQUENCE 

A/c hits abort height, 

ABORT TO HOVER lIEIGHT SEQUENCE 

Rel;1tive velocity > 5.5 

ft/sec. RUN FAST SEQUENCE 

A/C <3. ft of deck, CHASE SEQUtNCE 

A/C > 3.0 ft above deck, 

ABORT TO HOVER HEIGIIT 

EEQUENCE 

Ale < 3.0 ft above deck 

C!!ASE SEQUENCE 

A/C hits abort height, 

ABO~T TO SOVER HEIGHT SEQUENCE 

A/c <9.0 ft above deck, velocity 

of deck >4.5 ft/sec, ilL'N SEQUENCE 

Ale < CTNH, CHOP TIiROTTLE 

SEqUENCE 

Alc >3.0 ft above 'deck, 

. . 

ABORT TO HOVER REI CRT SEQUENCE 

Relat ive velocity ::> 5.5 ftl sec. 

RUN FAST SEqI!£NCE 

Relative velocity < 4~5 ftlsec, 

RUN SEQUENCE, till touch down 

r 

,.. 

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n 

p" 

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·.~-·I ~·.-,_.,l :==-:] ----., 

4---.. -_1 L_ .. _J r ____ •• i 

~ 

TIME HISTORY or SHIP AND Ale MOTION 

GROUP NO.1 

T/W !1AXI 1.05 

~ "" ~ '"' 

2-rCO-85 6:28:09' 

I NO. RUNSI I SEA STAT!:I 6 

ZH: 0.01 ·TAUENGI 0.30 KP: 0.001El 

[\·IJJ_ 

'\} AI:>CRAFT FLIGHT PATH 

o SHIP DECK MOTION , 

'.", 

f'. 

"" ~ 

" '. 

~ .- ... I'--' ~ 

---~ 

i-_._~.; 

1. 

1/'[\ t-. ~. ~ L 1\ ~1 

r\ il \ / \ / I~ t=:-:- -

/ 1\ I 1 II 1'\ . 

- - V \ V \ ,~ I . 1

1

- __ - - - ;~,I, J20.B2!l22.B 

TIME (SEC) 

1 Figure 31: Aircraft and Ship Motion Time History as Output from the Non-piloted -Simulation' 

L . 

·t5)ii~.:~~~::>~.:;tt;:ll.~~,·j;;:"·{';'i ;':~',:l;;r: >:l\<. ~"\!:;("?;~~·~.;5);'l'p.\,,,,,~2;, 7 ii"·i· ~,,'1/' 'f ~~~,!~~,. i';:'."'" to fA :~r;i.···-;·'> 

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TIME HI$TORY or SHIP AND Ale MOTION 

2·rEB·BS 6:41:20 

GROUP NO.1 NO. RUNS: 1 SEA STATE I 6 

T/H MAXI 1.05 ZWI 0.01 TAUENGI 0.30 KPI 0.0040 

: I ; ~I'~. . . J-~- ~~ I 1 1 ",. :' f;) -1- AIRCRAFT FLIGHT PATH 

o SHn' DECK HOTION o 

n 

~ 0 . --I- . " .' I 

:-1 I I I I I H """" •. 1-1-

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w· 

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~ ! \~ --'D VL I ,-I- lJ:o:,m~[jQ. g]62 .1lfi1.~mj ~_s '8. ~/O.Si'12. g]71. ~lfi. 'IJ1. .:000. 9m2. 9ilJi. SiB6. SlStJ. UJ90.~2.$91. ~6 .919I,.SJUO. 57 

Tl lE (SEC) 

t ___ . 

Figure 32: Aircraft and Ship Motion Time History as Output from the Non-Piloted Simulation 

~":-J 

I 

':'U!j 

~ ", 

I 

._-
,--..... 

0 

0 

0 

20.0 

SI~:utATIO!' 

Org. :loll-pilot ~Iod. Non-pilot. 15.0 

Piloted 

10.0 

TOUCH-DCAr.; 

VELOCITY. l\.o 

(ft/sec) 

5.0 

0.0 

1.0 

HUn) 

Sea State 6 

Velocity Coed. Sys. 

Z - -0.4 sec-1 

T " - 0.4 sec eng 

0 

0 0 

g 0 0 

9 0 0 

.J.. 

1.05 

TllRUST-TO-l-lEIG!lT RATIO, T/\i 

. 

e 

0 

1.10 I 

Figure 33: Comparison of the Influence of Thrust-to-weight Ratio 

on Touchdown Sink Rate--Augmented HUD Velocity Cornmann 

System of Piloted Simulation with Non-piloted System 

TEST 

CO~OITION 

o ORG CmID LOGIC 

o ~~D CmID LOGIC 

o PILOT HCD) V.C • 

• l'lLOT IIUD) A.C. 

HEA.'~ 

FLlGIIT 

TUIE 

(sec) 

Note: The o~ig1nal 

co~~nd logic was 

run from an initi~l 

altitude of 20 ft. 10.0L-----------------~ ____________ ~ ___ .J..~ 

1.00 1.OS 1.10 

TIIRUST-TO-WEIGHT RATIO, TN 

Figure 34: Comparison of the Influence of Thrust-to-\~eight Ratio 

on Flight Time--Aug~ented IillD with Velocity and 

Attitude Command Systems of Piloted Simulation and 

the Original and Hodified l-'light Path Command Logic 

of Non-piloted Simulation 

68 

..... -, 
;; . • #&il~?m*"e WmI5§A'M*%ii'C W' * • -Wi 

f"'"".. 

I 

\",\ ~ "-.(.~ 

.r 

--- --.. --~---- -..~- --.. .-.. - '- '-. :::-'"'---""'" 

A comparison of average touchdown velocities for a selected 

condition is shown in Figure 35. 

0 

0 

0 

a 

Z 

T 20.0 

SI:fULAT!ON ::,g 

HODIFIEO 0.3 

liON-PILOT 0.1 

PIWTED 0.3 

0.1 

1~.0 

--0.4 se,,-L w 

10.0 

TOUClI-Dmm 

VELOCITY, ~D 

(ft/sec) 

S.O 

"-

0.0 

1.0 

HUD3 

Sea State 6 

Velocity Comd. ,Sys. 

a 

a B 0 0 r.I 0 0 a 9 0 0 

0 0 

1.05 1.10 

TIlRUST-TO-WEIGlI'I' RATIO, TN 

Figure 35: Comparison of the Influence of Thrust-to-Weight Ratio 

and Engine Dynamics on Touchdown Sink Rate--

Augmented llUD \iith Velocity Command System and Modified 

Flight Path Command System. 

As can be seen the trendS are roughly correct for either 

command logic. However, there is still a fairly large bias that is 

unaccounted fe-:-. It should be noted that the data sho\ffi for the 

non-piloted simulation is a composite of data in which the pilot 

gains, aircraft vertical velocity damping, and variables mentioned 

in the T/W histogram section were being manipulated in an attempt to 

find optima (touchdown velocity and flight time) for each T/\~ 

ratio. Host of these runs were made shortly after the piloted 

simulations began, but before the data from the pilot'cd simulation 

were analy::cd. A more precise comparison requires another series of 

69 

.. " \ 

'-.. .. .. , 

1": 

. 

: ;:~,-~,t 

'~~:i~ '_ ~ ~.::.:~~) \ 

If{','·,' .' 

l~l!- 't·< i.' . 

iff~t~ -"~ 

-----.. -

':.~'.: . .:-; 
/ ~ 

. 

, 

J 

. i. ~.-"'-- , 

~-- " 

., { . :. .. 

non-piloted simulations to be run with variables similiar to those 

us"d in the piloted simulations. 

Initially, there was only interest in obtaining an indication 

from the non-piloted simulation as to whether or not the lower 

thrust-to-weight values were practical for landing in high sea 

states. In the early stag~s t~3 program was used to provide 

estimates of touchdown velocity means and standard deviation and the 

time required to la~d. As the work progres~ed it became clear that 

the fixed-base simulation output data used to construct thrust-toweight. ratio histograms, could then be used as another matching 

variable in determining the accuracy of the non-piloted simulation. 

There are 17 variables whose values determine in some way the 

shape of the histogram. These are: 

1. The initial starting altitude. 

2. The letdown time. 

3. The hover height altitude. 

4. The abort height altitude. 

5. The chop t.hrottle height. 

6-8. pilot gains. 

9. The pilot pure lag time. 

10. The pilot sidetask time. 

1 1. The ratio of 9. and 10. 

12. The pilot preception error noise amplitude. 

13. The pilot preception error noise frequency. 

14. The pilot internal noise amplitude. 

70 
>\ '.--' \ 

'-'\\ ",~-

,. 

I '. ' 

. .. ;.., -

\. 

15. The pilot internal noise frequency. 

16. Pilot lead time, time constant. 

17. Pilot lag time constant. 

Some of these non-piloteq simulation variables are easily fixed. 

Initial height, hover height, and abort heights are 'all displayed 

for the pilo~ and so are the same in both piloted and non-piloted 

simulations. Pilot internal noise amplitude was based on having 

each pilot,· and any other observers present, guess the velocity at 

which the pilot touched down as viewed on the head-up display 

monitors. The error was then determined and the noise standard 

deviation was set at 1.5 ft/sec after averaging over the number of' 

observers approximately 180 landings-. The res,t of the variables 

were set by makin9 an initial guess and then making 5 runs and 

looking to determin~ hbw the variable had chanojed the histogram. 

This process was repeated by either changing the variable again, or 

going to the next one. In this way a good match to the histogram 

was made for one of the cases {Figure 36}. It should be noted that 

there may be as much as 20-30\ change in ariy one point on the 

histogram for a given set of 5 runs (as can be seen in Figure 37). 

This is because 5 runs are not enough to get a good statistical 

representation, thus aggravating the difficulty of trying to obtain 

a good match. The non-piloted simulation showed that approximately 

40 runs (depending on the T/W ratio being used) is needed to p • .:>vide 

adequate statistics. .ihen histograms r€pr~senting only 5 runs are 

constructed based on 10 unknown variables, there is some difficulty 

in producing a good match. Another problem occurs when trying to 

71 

': \ . 

I 

",.'> 

.......

. ~i(.c~: 

·~~T·' 

',""'>-#:.."'" 

.. , •.... ," , -' ,"::-~ 

, \:~:: .. ;~. 

~;.,~ ... 

,..

. \;,;;,";' 

~":~I~ 
-.l 

N 

"\' 

) 

40.0 

30.0 

PERCENT 

OF 

Tare 20.0 

10.0 

0.0 

Figure 36: 

0.90 0.94 

) 

0.98 

6 Non-piloted simulstion 

o Piloted simulation 

1.02 

T/Wmax 

Z 

T

W 

eng 

• 1.05 

• -0.4 

.. 0.3 

1.06 

THRUST-TO-WEIOIT RATIO, T/W 

1.10 

Comparison of Thrust-to-Weight Use H1stogram~ as Obtained from the P~loted and 

Non-Piloted Simulations 

._) 

y------

;~]'s;i;q;::\:6{~;;,~;j· :,:;;,;:;.·~;J;.;j';;'&wi~t~\f~0~~~~\Y0j):.~~~~};'i~i?"\~%?;c;i:~;;;::"'i;';»:i;~I~?i::~¥;t '-, ..... , " 

.~ ~ . ,- -\ -\.:' , , \ i 

I. 

;.~--

\-

I 

----

" ..... '. 

--. 

~ .... / 

... -"'~-----

,>A 

.",," " 

( 

.\ 
r' 

·. ·--i 

50.0 r-..... ~--~ .......... ~ ..... ~ ..... ~ ..... ~ ..... ~ .......... ~ ..... __ ..... ~ 

40.0 

lO.O 

PERCENT 

OF 

TIME 

20.0 

10.0 

0.0 

0.92 

HtIDl 

Sea State 6 

Veloc'ity Cmd. Sys. 

T/W - 1.10 max 

Z - -0.4 sec-1 w 

T - O.l sec eng 

o Run No.s lll-315 

6 Run No.s 711-715 

0.98 1.04 

T!!RUST-TO-~ICRT RATIO. TN 

1.10 

Figure 37: Variation of Thrust~to-weight Ratio Use Histograms 

Due to a Statistically Small Number of Runs Used 

Per Data Point 

match the touchdown velocity and flight time averages for the same 

cases. It was found that once the major influence of variable 

changes on the histogram shape, was obtained, it was fairly easy to 

get a rough match of histograms. The Tf\i ratio for which the 

histogram peaks can be obtained through selection of the initial 

letdown velocity of the aircraft. The base can be broadened by 

select~ng higher values for the pilot gains. The sharpness of the 

I 

fillets between the base and peak were found to change somewhat ~ith 

the selection of pilot gains and the values used for frequency and 

magnitude in the pilot internal and preception noise models. The 

shape of the base; i.e., the number and size of tile peaks appeared to 

be mainly dependant on the ratio of three pilot gains used. It is 

73 

/ 

f 

. 

I 

i i, 

I 

I 

i 

I 

, \ 

.L 

"-"'~., 

,~.'."~:.~;' 

'" ," ~ .... '1 , 

1 • ,"'-;;. ~- " , " , 

, ',,' ,.~: 

- -Jl'F-,- ~ '. >: .. ~ 

'j' I, 
v , I 

, 

, .. , /: ~. / , , I , , I 

r-----'. 

also fairly easy to get a ·reasonable match of touchdown velocities 

and flight times. It is not easy to get all three of these results 

to give a reasonable match simultaneously. This may be due to 

s9mething inherent in either the piloting technique or aircraft 

model being flown on the fixed-base simulation that is not being 

modeled accurately in the non-piloted simulation, or that the right 

combination of values for the variables has not been found. The 

latter should be explored further using parameter identification 

techniques. It would ultimately be hoped that all three could be 

matched for a couple of case·s. and then the values of the variables 

be determined analytically to· propuce matching values for other 

cases. This pilot model would then be a good too.l for predicting 

the landing performance of any VTOL aircraft onto any type of ship. 

74 

\. 

\ 

\ : 

.... 

". 

/ 
I I 

f 1/ .I, " " f ' I / i 

I 

I 

I 

\, 

.' 

. ' 

V. CONCLUSIONS. 

The problem of determining the vertical axis control 

requirements for lariding a VToL aircraft on a moving ship deck in 

various sea states is examined. Both a fixed-bas~ piloted 

simulation and a non-piloted batch simulation were used to determine 

the landing ~rformance as influenced by thrust-to-weight ratio, 

vertical damping, and engine lags~ 

The piloted simulation was run using a fixed-base simulator at 

N.A.S.A.Ames Research Cent~r. Simplified versions of an existing 

AV-SA l~rrier model and an existing head-up display format were 

used. The ship model used was that of a DD963 class destroyer. Two 

pilots were used to obtain data and to give pilot ratings based on 

the Cooper-Harper pilot rating scale. 

A surprising result of this simulation was that, with a good 

station keeping control system and with statistical ship motion 

displayed on the head-up display, pilots could consistently ~rform 

safe landings in sea state 6, with handling qualities that were 

adequate at thrust-to-weight ratios greater than 1.03 and even 

marginally adequate down to thrust-to-weight ratios of 1.01. These 

results should hold quite generally provided that a thrust-to-weight 

ratio of 1 + ~ is interpreted as meaning that the pilot always has 

the capability of accelerating the aircraft at 6g upward even in the 

presence of ground effect and hot gas reingestion. 

Preliminary work with a non-piloted simulation showed thqt with 

a good strategy and th~ right information, a pilot should be able to 

75 

:' I 

.I , / 

, 

. ! , 

I 

i 

I 

• I 

I 

I 

I 

! 

\i 

, 

'\ !\ 

1 \ 

, 

i 

'f 

:~:;~;f : :·~·;.;---f 

"> 

::\:,,~-' 

>~.,~; 

:~.~-~- t 

I 

:~~~.:~; . 

\ 

I 

" ,

., 
..,. 

,. 

,..-... 

r 

• I /.~ /. 

.. .I.,f i· 

I 

I 

I 

land a VTOL type aircraft vertically aboard a 00963 class destroyer 

under sea state 6 conditions, in an adequately controllable manner, 

with thrust-to-weight ratios as low as 1.01, engine lags as high as 

0.7 sec, and vertical velocity damping of 0.2 sees, without 

exceeding a 12 ft/sec landing gear limit. This non-piloted 

simulation showed an overall average touchdown velocity of 5.8 

ft/sec and an average flight time of 32.5 seconds. Results were 

then obtained from a piloted fixed-base simulation in order to 

verify the non-piloted results. Simi liar results were obtained, 

wlth an average touchdown velocity of 6.7 ft/sec and flight time 

average of 33 seconds. In addition, the pilot ratings indicate. 

satisfa~tory (l~vel 1) handling qualities for sea state 6 conditons 

and thrust-to-weight ratios as low as 1.03 and adequate (level 2) 

handling qualities for thrust-to-weight ratios as low as 1.01. 

Pilot ratings showed the expected results of being more 

favorable as T/W ratio increased, up to the maximum tested of 1.1, 

and with increasing vertical velocity damping, up to the maximum of 

-0.4 sec-1 tested, and with the faster responding engine, engine lag 

of 0.3 sees. The pilots also demonstrated lower touchdown 

velocities when presented with the ship motion boundaries, and 

aircraft hover and abort chase height lines on the head-up display, 

then when presented only with the ship deck position symbol. The 

simulation also showed that the pilot was capable of obtaining 

similiar results when flying either the tran3lational velocity 

command system or attitude command system. pilot ratings indicate 

76 

\, 

\ 

t;: 

i 

,I 

... j 
! 

~ 

f 

/ " 

/ I /l···· ; 

/ 

i 

that the translational velocity command system was perferred 

however. 

Based on the current ncn-piloted simulation, it is believed· 

that an extension can be made to determine pilo.ted results for other 

T/W ratios, engine lags, vertical velocity damping, and ship 

classes, under various sea state conditions. This is based on the 

assumption that, using parameter Indentification tecihniques, touchdown velocities, flight times, and T/W use histograms can be made to 

match the current piloted simulation data·. 

Although the simulation indicates that aircraft can be landeri 

vertically at much lower T/W ratios than previously suspected, even 

with the positioning siaetask, it remains to be seen how well low 

thrust-to-weight ratios will work when a full range of side tasks 

inherent in flying actual aircraft are employed. In addition, more 

wot:k should be done to dete.rmine the effects on minimum T/W· ratio of 

non-linear elements which were not examined in this simulation. The 

effects of suckdown, fountains, hot gas ingestion, and height 

dependant mean winds can all significantly effect minimum T/W. 

As a practical point regarding the simulation, additional 

research would have to be conducted to ascertain how well 

instrumentation aboard ship can determine the mean ship position and 

the motion boundaries, how long in advance of the aircraft arrival 

the motion would have to be monitored to provide accurate results, 

and h~~ well the aircraft/ship systems can determine real time deck 

position. 

77 

I • I 

, 

I 

. \ 

! 

t 

i 

I 

, 

, 

I 

I 

.;/'" 

, 

/. 

..... 
II 1-

Ref. No. 

1 • 

2. 

3. 

4. 

S. 

6. 

7. 

8. 

9. 

10. 

/ / 

IV. REFERENCES 

.<eference 

Nave, Ronald L., "Analysis of Height Control Power 

Requirements for V/STOL Aircraft," NADC Report No. 

79211-60, July 30, 1979. 

polk, Clyde H., and Phater, Anil V., "Evaluation of ~_ 

Real Time Prediction Guidance Law for Landing VTOL 

Aircraft at Sea," AlAA 84-2673-CP, December 1984. 

Baitis, A. E., "The Influence of Ship Motions on 

Operations of SH-2F Helicopters from OE-10S2~Class 

Ships: Sea Trial l'ii th USS Bowen (OE-l 079) , " 

Report No. SPO-SS6-01, July ,-975. 

Sider, H. H., and Ooolin, B •. F., "On The Feasibility 

of Real-Time Prediction of Aircraft Carrier Hotion at 

Sea," IEEE Transactioml On Automatic Control, Vol. AC28, No.3, March 1983. 

McHuldroch, C. G., "VTOL Controls .for Shipboard Landing," 

Unlversity of Victoria, Canada, 1976. 

Nave, R. L., "A Computerized VSTOL/Small Platform 

Landing Dynamics Investigatl-on Model," N1\PC-77024-30, 

1977 • 

o. T. McRuer, o. Gr:aham, E. Kredel, "Manual Control of 

Single Loop System," Journal of the Franklin Institute 

Volume 283, No.1, January 1967. 

Battis, A. E., Meyers, W. G., and Applebee T. R., "A 

Non-Aviation ship Motion Data Base for the 00963, CG 

26, FF 1052, FFG 7 and the FF 1040 Ship Classes," 

OTNSRDC Report No. SPD-738-01, December 1976. 

Fortenbaugh, R. L., "Progress in Mathematical Modeling 

of the-Aircraft Operational Environment of 00963 Class 

Ships," AIAA paper No. 79-1677, August 6-8, 1979. 

Merrick, Vernon K., "Simulatiol'l Evaluation of Two VTOL 

Control/Display Systems in WC Approach and Shipboard 

Landing," NASA TM 85996, December 1984. 

78 

. ",.':. 

, f~1 ':, i .~.~~. ; , 

\ 

~. ~ " 

1-." 

. 

- 
Ref. No. 

11. 

12. 

13. 

REFERENCES, continued 

Reference 

Ferris, G. G., Merrick, V. K., and Gerdes, R. M., 

"Simulation EValuation of Flight Controls and Display 

Concepts for VTOL Shipboard Operations," AlAA paper 

No. 83-2173, Augu~t 1983. 

Cooper, G. E., and Harper, R. P. Jr., ".The Use of 

pilot Rating in the Evaluation of Aircraft Handling 

QUalities," NASA TN D-5153, April 1969. 

Roskam, J., -Airplane Flight Dynamics and Autcmatic 

Flight conttols,"Roskam Aviation and Engineering 

Corporation, 1975. 

79 

"'-. 

I '" •. ~ 

.':;";y-

,:\" 

'",,:' 

! 

/ 

I 

/ 

'. . 

\ 

/ 

/ 
***********************~** 

* 

* 

* 

.VII. APPENDIX A 

* 

* 

* 

************************.* 

A1 

" , 

, \ " 

\. 

, 

I 
, 

-\ -~~ I' --<I .' " .. f *. ' . ." ~ . . ' . . ~ I :: h' '. "0 ___ •• _______ _ 

The computer program for the' no'n-piloted simulation was 

discussed in general in the main bo~y of the report. Some more 

detailed information regarding certain aspects' of the program are 

described belo-". 

A. A~ROXIMATE INVERSE LAPLACE TRANSFORM METHOD. 

The numerical iteration methoa used for solving the 

'differential equations used in the computer program is the 

Approximate Invers~ Laplace (A.I.L.) method (Reference 13). It is 

assumed that th~ equations fo~ the transfer functions are in the 

form: 

where m > nand Bm is not equal to ,zero. 

Then using the following recurrence formula: 

1 

B 

m 

m-1 

[A. - L B.C. .J n-~+1 j=O J ~-m+J 

where A x B 

x o for x < 0, and C x o for x .. o. 

Xes) can be rewritten as 

xes) + ••• 

The inverse transform of equation (A3) is 

A2 

(A2) 

(A3) 
. ". , / , i..... "i-'· . '/ 

~/"' 

.... _._ I .. 

, I 

, 

i 

.. 

' . 

(A4) 

The derivatives can then be found from differentiation: 

(AS) 

(A6) 

The A.t.L. method ~iloW5 the use of time varying coefficients 

in the transfer functio~. The method is based on the Assumption 

that a time interval, 6t, can be found during which all of the 

coefficients can be considered as constant. A set of points 

calculated in one interval of time ~s then used as tIle initial 

condi tions for the next calculation, and the prc.·cesp. i9 repeated. 

Hare information on the A.t.L. method may be obtained from 

Reference 13. 

B. A. I .• L. I14Pf.EHENTATION 

~ne specific equations ~qed for the non-piloted simulation were 

derived as follows: 

51 ..... dH~J KP,! TL s + 1) l Ht_"" g/ELC z .... 

A) 

~-~, -l (TG s + 1) I 

A 

dH '" KP(TL R + " 

(TG s + 1) 

Can be expressed a5 

- 3 2 ... s + (l/ELC + Z )9 + Zw/ELC s w 

B 

lit, . 

A3 

~---~ 

:$"~~~~~~ 
Then 

• 1 Y + TG Y - dH 

Using Laplace trans,formation, 

y • r 

• 

y .. sr - yeO) 

Therefore, 

or 

Sr_y(0)+_1_ r TG 

r _ y(O)s + dH: 

2 

s + 1fTG s 

dR zaS 

using' the A.I.L. method: 

1\0 .. 1 

A1 .. yeO) 

BO .. 0.0 

B1 .. 1/TG 

B2 = 1.0 

And from the reversion forumula: 

C1 = A1 

C2 .. AO -81 * Cl 

C2A = AD 

C3l\ .. -81 • C2A 

1\4 

(A1) 

(AB) 

Where dH is considered 

a constant for the time 

interval under 

consideration. (A9) 

(dH will be reintroduced 

at a later point) 

(A10) 

(A 11) 

C2B = -Bl * C1 

C38 = -81 * C2B 

, , . : .. 
B) 

C4A .. -B1 • C3A C4B • -B\ * C3B 

• • 

• 

Then 

yA ~ C, + C2A * IT' + C3A * IT2 + ••• 

'VB- C2B • IT' + C3B • IT2 + ••• 

• YA = C2A + C3A * IT1 + C4A • IT2 + ••• 

• yB a C2B + C3B * IT1 + C4B * IT2 + ••• 

For- the next. iteration 

And 

Or 

y(O) &: y n n-1 

Y(O) n 

y~ ~ [(YA * dH + YB) + TL' (yA * dH + yell 'l'G 

Ht:,. q/ELC ... 3 2 s + (1/ELC + Z )s + Z.jEIJ: s w 

... 

Zw • z + (1/ELC + Z ):: + -- Z '" ~ Ht 

w ELC ELC 

AS 

Ht 

z 

" 

(A13) 

(A14) 

(A1S) 

(1\16) 

I. 

;... 

-- ---- ----::---,~ -----.-~--- ... &_.~~ ........ "-~-
r 

/---

substituting (11.15) for Ht: 

Z 1 ,.,. 9 TL· • z + (EiC + z,.,)z + ELC z - ELC{KP TG (YA * dH + YB) 

+ ~L (yA * dH + yB)]} (A17) 

But dH '" S1 -ZI 

B 

Therefore, (11.17) can be re,.,itten as I 

Z I r 

• Z + (1/ELC ,., +~ {KP TL • 1 + Z )z + --z [yA +- (yA)]} Z ,., ELC ELC TG TL 

a_CL- h<p ~ [(rA . + ~L (yA ELC * S1 + yB) , 

I 

A 

Or 

1 Z,., . Z + (ELC + Z )z +-- Z + D A ,., ELC 

Using Laplace transform: 

z .. 1; 

Z .. S1; - z(C; 

2 . Z = 9 1; - 5Z(0) - z(O) 

3 2 . Z .. 9 1; - s z(0) - 9Z(O) - z(O) 

And therefore 

3 s 1; - .2 . s z(O) - Sz(O) - z(O) 

Z.., 

+ -- [S I; - z ( 0 ») + BI; ELC 

1 

+ (-- + Z )[5 ELC ,., 

=.-5 

1\6 

* S1 .. yB)]} (11.18) 

----l 

(A19) 

2 . I; - 5Z(O) - z (0)] + 

(1\21 ) 

, , . , ~ 

- .. 
where" B is conside~ed ~.coAstant for the time 

interval under consideration. 

Or 

3 1 • 2 

z(O)s + [EtC + Zw)z(O) + ~(O)19 

t - --------------------~---------- 4 1 3 Zw 2 

s + (ELC + Zw)s + ELC S + Bs 

Zw 1 • [ELC' z(O) + (- + Z )z(O) + z(O)]s + A 

+ ~~ _________ E~LC~ __ ~w ____ ~.------------

4' 3 Zw 2 s + (iLC + Zw)s + ELC' S + Bs 

Using the A.I.L. method: 

MO D A 

Z 

M' ~ ~ z(O) + (--'- + zw)ieo) + zeO) ELC ELC 

1 . AA2 .. (--+ Z )zeO) + z(a) ELC w 

M3 .. z(a) 

BO .. 0.0 

81 = n 

Z 

B2 w = --- ELC 

R3 = 1 (- + Z ) ELC w 

B4 .. 1.0 

And from the recursion fo~mula: 

1 ~ 

C i .. -4 [ 1\4 . - r 8 . C ( • .) 4) R -1 j::O J "-J-

(1\22) 

(1\23) 

(1\24) 

I 

I 

~, 
r 

ct .. A3 

C2 a A2 - B3 • C1 

C3 • At - B2 * C1 - B3 * C2 

C4 .. AO - B1 * ct - B2 * C2 - B3 * C3 

cs = -B1 * C2 - B2 * C3 - B3 * C4 (A2S) 

• 

Then, 

Z .. C1 + C2 * IT1 + C3 • IT2 + 

Z .. C2 + C3 * IT1 + C4 * IT2 + 

Z = C3 + C4 * IT1 + CS * iT2 + (A26) 

For the n~xt iteration, 

z(O)n ft zn_1 

. 

z(O) .;. z n n-1 

z(O) = z n n-1 (A27) 

A block-type diagram showing how the A.I.L. equations are 

implemented in the-computer program for the pilot servo and aircra-ft 

transfer functions is shown in Figure A1. 

A series of runs wore made to determine the number of h.I.L. 

coefficients and size of time step required to obtain accurate 

results and a stable program. It was found that 6 terms and a time 

ste? of 0 •. 01 seconds provided good Iesults. 

AS 

, . 

, 
/' 

") ) ') 

.. '.",,~~ .~ -~ .... 1"'::--' -;-,/,,-

Plt.OT BI.OCK I A I RCRA IT BT,OCK 

I 

I 

I 

I 

I 

I 

I 

I 

I 

IItA I 

)0 I I rt,1 + ~'l.!!J rlf' AI 10 I [I ,'I ,Yoj .. ' ".,fo,,) §J + 1 ,rojl<, f ,:~ G~ol +( ,~c +') I, ~ol 

I 

51 L~.n drt.! ~I Al+t., +:1 ~YDO ~ 

Figure Al: Block Diagram of the A.I.L. Impl~mentation in ,the Non-piloted Simulation 

~;.; -- ----'--',------
c. DEVELOPMENT OF 'rut: SHIP MoTION HODEL 

. ~ ~'. 

. , 

To obtain the amplitudes for. the individual ~inusoidal 

components, the following method was employed. First, the twoparameter Bretschneider wave spectrum is used to obtain the wave 

spectrum: 

2 483.5 H 

. 4' 

-1944.s/(w T ) e 0 -s Sw (III) .. ------='----::4--::5--:------

To III 

This is transformed 'using the following relationship: 

2 

III V cos( \l ) S S III = III - ~-~--~- e 9 

(A2BJ 

(A29) 

To obtain the effective frequency based on ship velocity and heading 

relative to the wave'S. The ship motion spec;t:rum can then be 

obtained as follows. 

0111 ~ii(1II ) = S (III ) -6- RAO. (III ) e we III :I.e e 

(A30) 

A computer program, SHPREF.FOR, was developed to accept a table of 

RAO values corLesponding to a set of values, and then store the 

information in a data file. This table is'then read hy the program, 

and the above calculations are made. A numarical table and a plot 

of ¢ versus lIIe is then output as the finished result; see Figures A2 

and A3. ~le plot was then sectioned by hand into six cOmponents and 

the amplitudes for these six frequencies were determined by the 

following relationshir: 

A10 

I 

I {,.,. ' 

j l8, 
r\ n 

IT 

f] 

p , I 

.. 

11 

I ! 

; i 

r 

! 

.. 

-.... 

t..t:. 0.215~281 

I . 0.274"::62E 

\.iE- 0.';~oj9382 

!JE- 0.3975546 

wE- 0.4621124 

WE- 0.5286109 

WE- 0.5970S05 

wE- 0.667<1311 

wE- 0.7397528 

WE- 0.776641<1 

",E- 0.8140153 

WE- 0.8518746 

wE- 0.8902190 

WE- 0.9290487 

WE- 0.9683636 

WE- 1.048449 

wE- 1.130476 

'. wE- 1.214444 

WE- 1.300352 

WE- 1.388202 

WE- 1.477993 

WE- 1.569724 

t..E- 1.663397 

WE- 1.759011 

ORlU!NAL r=.'i ::::! 15 

C"- POOR QUAUi"l 

i SHIP HOTION SPECTRA 

. , 

:~--~--~~H-~--~~~-+~----+---~ ,. 

~,~~~~~~~~.~~I.~i~ 0.' 0.' o.a t.' 1.0 1.1 !., I.' lit 'IftIr.a:I 

.~ ... - PHI- O.OOOOOOOE+OO St..(l )- O.OOOOOOOE+OO 

PHI- 1.5390024E-15 51.1 ( I>- 1.4029235E-14 

FI'H- 4.5292235E-07 51.1(1)- 4.1409426E.-06 

PhI- 1.31302HE-03 SW( 1>- 1.2105345E-02 

PHr- 4.6292055E-02 SW(!)- 0.'B41457 

PhI- 0.238097<1 51./( I)- 2.309751 

PHI- 0'.5212409 SW<I )- <1.963209 

PMI- 0.7592450 SW<I) - 6.730645 

PHI- 0.9081539 SW(U- 7.144627 

PHI- 0.9543847 SW<I )- 6.967761 

PHI- 0 .. 9798968 St.(I)- 6.630970 

PHI- 0.9986890 St.;(I)· . 6.196284 

PHI- 1.000000 5t.i( 1>. 5.711943 

PHI- 0.9313180 5w( I>- 5.212773 

PHI- 0.7695953 St.i( 1>- 4.722323 

PHI- 0.3381195 5101(1)- 3.820434 

PHI- 8.9954354E-02 5101(1)- 3.059384 

PHI- 1.44I:l<lOI1E-02 51.)( 1>- 2.4<11777 

PHI- 1.2603654E-03 S/.,( 1>- 1.950421 

PHI- 1.7391144£-0<1 SW( 1>- 1.563076 

PHI- 3.7044092E-04 SW(f)- 1. ~58615 

F-HI- 2.83058211:.-04 SW( 1>- 1.019108 

PltI- 1.3718315E-04 51.1( I>- 0.8301120 

PHI- 5.8686554E-05 510(1)- 0.06803094 

PHII"Ar- 7.771437 »~ 6I.~ ML"S- 2.094<100 115- 6.900000 TOI 

Figure A2: EXample output from SHPREF.FOR with Plot of 

Ship Motion S.<;>ecturm for the 00962 in Heave 

All 

10.60000 

10

~ 
r 

.! 

""r' 

.~ , 

't " •• '-. 

r 

.1.1 o.a 

ta.E- 0.2155281 PHI- ~IC_ 0.2742626 PHI- ~ . 0.3349382 PHIWE- 0.3975548 PHIwE- 0.4621124 PHI';' 

WE- 0.5286109 PHIWE- 0.5970505 PHI- . 

"'E- 0.6674311 PHI- I~E~ 0.7397528 PHIWE- 0.1766'1H PHIWE- 0.8140153 PHIWE- 0.85187'16 PH!-

t.;E- 0.8902190 PHIWE- 0.9290487 PHIIoIE- 0.9683636 PHIWE- 1.048449 PHIJ.;E- 1.130476 PHIIoIE- 1.214'144 PHI-

,"E- 1.300352 PHIWE- 1.388202 PHI101[- 1.477993 PHIWE- 1.569724 PHIWE- 1.663397 PHIWE- 1.759011 PHISHiP MOTION SP~CTRA 

I I 

I • ;11: ~ I,. t.' 1.0 1.2 It: llUIIJ"...a:J 

O.OOOOOOOE+OO SW<I }-

1.0081215E-H:i SW(~}-

4. 12S'0222E-08 5W( I>-

2.2123220E-04 Sloi(l}- 1.261£537~-02 5W(I}" 0.1025751 SW(l)-

0.3154627 SW<I }-

0.5932509 5W<I )-

0.8'116853 5loi(I)-

0.9302165 51oi<I}- 0.9839358 5101(1)-

1.000000 SW( I>-

0.9649336 SW<I )-

0.8602818 SW( 1>-

0.6967936 SW( 1>-

0.3589893 SW(l}-

0.1623948 SW( 0-

6.8711556E-02 SIoi<I )-

2.56937221:-02 SW<I )-

8. 2251765E -03 5W(I)-

1.22601621::-03 51.1( I)-

8.9359164E-05 5W(l)-

7.98591S7E-05 SI.oI(I)-

7.7392666E-05 51.0/( 1>-

/.I I.' 

O.OOOOOOOE+OO • 

1.4029235E-14 

'1. 1 '10942GE -06 

1.2105345E-02 

0.4341457 

2.309751 

"1.963209 

6.730645 

7.1'1'1627 

6.907761 

6.630970 

6.196284 

5.711943 

5.212773 

'1.722323 

3.820434 

3.059384 

2.441777 

1.950421 

1.563076 

1.258616 

1.019108 

0.8301120 

0.6803094 

PHI MAX- 0.5683722 MUS- 2.09<1400 H5- 6.900000 TOGD963P.OAT 

'. 

Figure 1\3: EXample OUtput from SHPREF.FOR with Plot of 

Ship Motion Spectrum for the 00962 in Pitch 

1\12 

10.60000 

I 

! 

I 

I 

I 

I 

j 

I 

i 

"" 
I, ; 

I I 

t . 

L. 

[',' 

, 

; . .\ 

E 

i,1 

II' 

Ll. 

(A31) 

A component of the ship motion could then be calculated as: 

6 

i(t) = E An COS(We t - ~ii + En) 

na l n 

(A32) 

where ~ii is directly available from the ship data base information 

contained in Reference 8. En is a random phase angle which, is 

calculated from a random number generator with the output scaled to 

give values· between· 0 and 6.242 radians. 

In the computer-run simulation,the ship motion is calculated in 

the GENeral Ship' ~Iotio'l, GENSM.FOR, subroutine. This subroutine 

uses th~ associated amplitudes and phase angles corresponding to a 

particular sea state as stored in th.e data file, GENSM.DAT. 

D. TURBULENCE HODEL. 

The turbulence model for the computer-run simulation is located 

in the subroutine TURB3.FOR. It consists of a random number 

generator and the following principle equations: 

. n (1/TC' / 2 ~= W ~+ C1 (2w ) ) n v n 

. A = AR CA33) R + TC AR 

The bandwidth, wn' was obtained from page 10, Figure 15, of 

Reference 9. Tl1e sigma values were obtained from a strip-chart 

recording of the acceleration of the vertical axis for the AVB-B 

A13 
• 

during a fixed-base simulation, as presented in Figure 11.4. The 

(1/TC ** 0.5) term is a correction for the effects of usin~ digital 

computation. The random number inp~t ~s represented by n. Becuase 

this is a filter acting to shape white-noise, it is neces'sary to 

precycle it initiully ~fore inpueing values into the simulation. 

Tlais is also accomplished as part of t.he subroutine. 

E. HOVER HEIGHT INPUT 

The ~'.lbroutine IN4.FOR provides the values for the flight path 

command logic hover height. The sequenoe starts at an initial 

altitude and, follows a shc;tllow cosine path to a selected hover 

f~ altitude as a function of time. The hOVer altitude is then , 

maintained as a constant fer the remainder of the run. It is also 

possible to configure tl1e subroutine to change to a second hover 

height during the run, although this function was not used other 

than for initial testing of pilot lead and lag time constants. 

Variables which can be adjusted before running a simulation are, the 

initial altitude, AMPA + AMPB, the hover altitude, AMPA, and the i ' 

' .. rate of descent, through the frequency term of the cosin~ function, 

WNS. 

F. DATA PL9TTING 

The subroutine HPLT2.FOR providC:!s the instruction to the'system 

library and DISSPLA software to produce a plot of the desired data 

11.14 
) 

»I 

~ 

VI 

')7-

,~ -j 

..... '.~' '. 

'ligure M: Strip Chart Recording of vertical Acceleration 

Due to Turbulence as Obtained from the AV-SA 

Fixed-Base Simulation Facilities and Used to 

Determine Magnitude of Turbulence V.alues fo~ 

the TUrbulence Modeling in the ~on-Piloted 

Si~ulat~on ' 

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

-... -..... ,." ·r'··----··· 

00 

."''''' ~ '7.J !i~ 02 

O~· 

~r-

,o"Tl c: .... );;It:. 

r- ,.1 

:J ul 

") 

_/, ." 

( 
· '. 

\ '.' 

~ to ' 

I . ~. 

( 

.,:' , \ '., 

&t the end of program execution. Currently, plots of tha T/W ratio 

used as a percentage of time, and statistical variations as a 

function of the number of runs made are output (for examples see 

Figures A5-A7). A plot ~f the aircraft fliqhtpath and ship deck 

position time history can also be set up with small changes in 

program configuration (for examples see .Figures 31 and 32). A 

tabular listing of the runs made for each test configuration is also 

output, an example is given in Figure A8. 

The following table is' a listing of variable names and their 

approprii!te values which must be,dited into the indicated programs 

when a change in sea state is made. 

Table A1: Valued for Variable Which Must Be Edited 

into Prog.ram with Chan'C]e in Sea State 

VARIABLES W1UGlI API. FUNCTION OF StA STATE 

Subroutine Varuble Sea State 

~lame 0 4 5 6 

TlV.NO.FOll \.'N 2.70 6.15 7.79 7.79 

SICiA 0.001 0.007 0.01 0.01 

I114.FOR A.''J'A 0.0 5.0 7.0 9.0 

AMPS 40.0 35.0 33.0 31.0 

VPAlC ABRntT 0.0 1.5 3.0 6.0 

C~SM.DAt 

(Version tlo. - - ;4 ;5 ;6 

VASCON.DAr Supplies the input data for the test conditions 

and 1s also changed 48 apporprlate. 

1\16 

--------------------
') 

;::.0 

.... 

-..I 

1 

~r:-" . •. " . 

..., .......... . :;;.'~ ... 

w 

~ 

..... 

I--

... 

o 

..., 

o 

LON 

1-- 0 

13 

CJ 

a::: 

w 

0... 

I 

o 

c:; 

o 

• 

. 

I 

~ 

\ 

) 

T/H RATIO STATISTICS 

yf 

/ 

J~ ~ 

~ 

.-r---

-f-

.L">. 

ero.f.2 0.84 0.86 0.8e 0.90 0.92 0.94 0.95 0.58 1.00 1.02 1.01 1.05 l.OB 1.10 1.12 I.Ii 1.16 I.la 1.2Il 

T/H RATIO 

Figure AS: Example plot of Thrust-to-weight Ratio Use as Output from the Non-piloted Simulation 

'r''''~---' .. 

\ 

\, 

) 
, ! , I 

.,I 

• 

·1 

'., \' 

'" 

.. ~ 

~1>J : ' --'L 

\ ,! 

:: 

CD 

,.1 _ . . ;., .. ,,~ 

" r--:-:--

C 

tfl 

"" 

u 

We '" . "'0 

!;:"" 

u W 

'" 

0 

W~ 0 

:::J 

!-< 

..... 

Z 

Cl 

0: 0 ;C. 0 

C 

Ln 

o 

c 

-) 

STATISTICAL TRENDS IN V.A.S. SIMULATIONS 

,/"v --- ...... 

j 

~~ "-- --

0.0 

~ 

~ I-- -...-

Ii] FLICIIT Tum MEAN 

---

0 FLICHT TUIE VARIENCE 

0 TOUCII-OOI."N VELOCITY ME:AN 

0 TOUCH-DOWN VELOCITY VARIENCE 

./ 

./ 

25.0 

NUMBER or RUNS 

.. 

I 

i 

SO.O 

Figure A6: Example Plot Showing the Variation in Variance and Means for Flight Time and Touchdown 

__ ve~~citL~~ Influenced by the Humber of Runs Made _~~ the Non-piloted Simulation 

....... .,. ~.'" _ ..... _'t ___ ... - - - . . .. ----. ---........ ~.- . f:~', 

') 

~ 
\ 

J! 

:I>' 

\0 

. . '.i~~~i ... 

- ,:I ----. •...... -. 

~ 

-.'.. , , -. 

.. '" 

) 

STATISTICAL TRENOS IN V.A.S. SIMULATIONS 

u .. .•• ,- --r------ -----------,-

! 

I 

I 

a 

o 

'2 

SHIP DECK POSITION VARIENCE 

snIP DECK VELOCITY VARItNCE 

snIP DECK POSITION MEAN 

o SHIP DECK VELOCITY MEAN c 

'" 1-----------1· -,----

i~l~~-F_~~=b~-=:P~~~~==~k ~ 1 , 

I 

I 

I 

~~ .. ?-1~~~~{----g--=--~-~:> 

, 

! 

~J- 1 .. - -,- I I I 1-

lUI 25.0 50.0 75.0 100.0 125.0 150.0 175.U 

iI NUMBER or RUNS 

Figure A7: EXample Plot Showing the Variation in Variance and Means for Ship Motion as Influenced 

by the Number of Runs Made, Where Number of Runs is Seconds in This Ca~e 

---------------

--/"'. .... -- ...... _--.. __ .-.-.--------_ ... _-_.--....... -........ . 

.. _) 

. ~'", 

-~~ 

I· 
,I""""' 

PRi::i:NAL F,i.'~::; IS 

PE POOR QUALITY ._ •• 4 •••••••••••••••••• ~ 

( :LF :-or.: 

'~L· 0.15uoooo ELe- U.7000000 i"'- 1.010000 z",· u.l000000 SE':: S.- 6 KP- '1. 0000002E -0:: 

""' ....... ---" .............. -

RLN LT. F. T. T[\tEL ·i"IliRD TH~EA'" T~I"AX Thl"II't ----- ------ ------- ------- -------- ------- ------- 1 21.:65.7 21.2 r;. " 7.S 0.'=68 1.0tO 0.877 ... 

.. 121.'1 24.1 11.4 11.': .O.SSE 1.010 0.950 

3 ~O~€.6 Z4.2 S.8 S.l O.H:l 1.010 0.';50 

4 384.2 23.4 3.8 E.S 0.S85 1.010 o.ssa . , 5 2524.8 lS.7 B.8 ·10.13 0.S67 1.010 O.SSO 

/ 6 774.4 17 .13 6.S G.4 0.SG7 1.010 0.577 

7 ~825.2 27.0 3.2 H .w q a.see I.OtO 0.950 

S 262.6 22.2 5.1 c _.<;; c 0.9'3E 1. 01 a 0.950 

9 2'12£.3 22.5 6.7 6.7 O.St'S 1.010 0.950 

10 249i .'3 2~.8 6.9 9.7 0.965 1;':'110 0.£50 

11 2404.7 11. 0. 5.E 10.7 0.98'; ~ .010 0.577 

12 545.4 24.9 a. 'I S.,; 0.967 1.010 0.950 

13 1816.2 19.4 0.4 11. 1 o.ges 1.010 0.950 14 672.8 22.3 3.4 0.5 0.990 1.010 0.950 15 3490.0 2'1.6 6.9 9.9 0.986 1.010 0.950 16 962.4 16.9 3.2 8.6 0.587 1.010 0.977 '7 3467.7 1'0 ~.l. ... 4.2 6.5 0.£38 1.010 0.577 

.<1 543.1 18.9 to.5 11.8 0.986 LOW 0.950 

r~ 19 3075.1 24.5 4.6 9.3 0.991 1.010 O~S77 20 3225.8 20.7 10.2 10.2 0.9a5 ~ .Oi 0 0.950 

21 2802.3 21.1 8.2 8.7 0.984 1.010 0.950 

22 1931.2 12.0 3.1 9.5 0.994 1.010 0.950 

23 1901.3 27.4 11.0 11.0 0.983 1.010 0.950 

24 1974.7 24.2 4.8 11.7 0.991 1.010 0.950 

25 2514.5 22.6 3.5 6.3 0.989 1.010 0.977 i 26 225~.:: 12.3 6.2 9.0 0.986 1.010 O.S77 I' 27 3116.8 19.9 5.5 7.6 (l.sea 1.010 0.977 ! 28 1216.4 30.6 2.3 9.0 0.9S:; 1.010 0.950 I 

29 129S.2 1£.1 0.6 7.4 0.£88 1.010 0.950 l 

, 30 2666.2 17.4 9.6 10.6 0.9&7 1.010 0.977 

31 2999.0 24.'1 3.7 7.3 0.~a6 1.010 0.950 

32 1157. a 23.2 3.3 5.0 0.9SS 1. 01 0 0.977 

33 165.6 12.4 3.5 S.5 0.S87 1.010 0.977 

3'1 2190.3 8.3 8.6 14.6 ' 0.990 1.010 0.978 .-, . 35 866.7 20.2 5.0' 9.a 0.984 1.010 0.950 I ~~ ... 36 3417.2 28.2 2.3 E.9 0.9£2 1 ;010 0.950 ; 'l 

37 2£84.7 17.2 5.2 8.5 0.S88 1.010 0.S50. 

38 615.9 16.1 6.5 9.0 0.9Se 1. 01 a t),S77 39 521.1 25.8 2.6 12.1 0.990 1.010 O.Sso 

40 2124.9 33.2 10.~ 10.3 0.9S1 1 .Cl a 0.£50 

41 1051.6 20.5 3.1 5.9 0.989 1.010 0.950 

12 165.8 12.2 2.8 9.5 0.9S7 1 • ()1 a 0.977 

f 1 3314.5 33.4 10.9 11.5 0.991 1. 01 0 0.950 .,4 2679.8 24.7 0.7 10.4 0.S89 1.010 0.950 

1, 45 2616.7 27.7 8.6 8.6 0.9a7 1.010 0.950 

, i 46 2302.2 25.9 1.4 10.6 0.p-a9 1.010 0.950 47 2956.3 21.2 7.7 7.7 0.984 1.010 0.950 

L_ 

Figure A8: Example Output of Run Conditions and Results 

for the Non-Piloted Simulation 

. ..J.-'" A20 

" , 

",. 

~.' 
r 

\ 

V 

-l8" 1083.0 16,0 3.4 8.6 0.986 1.010 0.977 

49 178-L8 27;9 4.:? 6.& 0.9S1 1.010 0.S50 

50 2Z04.6 '\':) '1 c.. ..... , 1.1 10.4 0.968 1.010 0·.950 .......... --- t I"EP~(SEC) : 21.55800 FT SICHSEC): 5.464322 

FTMAX: 33.40000 FTMI!';: 8.300000 

TeVEL MEAh(FT/S): 5.5~6001 TOVEL SIG(FT/S): 3.060933 

TOVELMAX: 11.40000 TOVELMIN: 0.4000000 

MTGUT MEAN: 0.9879200 I"TGUT SIG: 2.3633221E-03 

MTCUTMAX: 0.9940000 MTOUTMIN: 0.9840000 

TIME:l1:1':54 DATE: 4-JA~-85 

•••••••••••••••••••••••••••••••••••••••••• 

ftJRTRflN STO" 

Figure AS, continued: Example Output of Run Conditions and Results 

for the Non-piloted Simulation 

",;. A21 

(' 

I 

t· , ,"",,:,: 

. .~ 
r 

/' 

/' 

r---' 

; I 

. ' 

" , 

n 

n 

11 

I ! 

Ii 

1I 

[I 

/ 

G. LISTING OF NON-PILOTED SIMULATION COMPUTER 

VARIABLE DEFINITIONS AND PROGRAM LISTINGS. 

.. ' 

1\22 

... ---...~ 

<II 

j 

I 

I 

I' 

1. 

! 

j .' I I: 

I' 

.' . 

'';'; 

,.,. 

· I 

• I · 

; , 

\, 

.. 

\: 
- . r" n : 

[j 

U 

U 

! 

r-" 

I. 

i 

[ 

....... r 

I-,-

/_ .... 

--------~~-. ,.-, 

LIST OF VARIABLES FOR VPA1C.FOR PROGRAM 

NAME DESRIPTION 

AO CONSTANT FOR TijE ZEROTH ORDER NUMERATOR TERM OF THE PILOT 

TRANSFER rJNCTION. 

A 1 CONSTANT FOR THE FIRST ORD.ER NUMERATOR TERM OF THE PILOT 

TRANSFER FUNCTION. 

AAO CONSTANT FOR THE ZEROTH ORDER NUMERATOR TERM OF THE AIRCRAFT TRANSFER FUNCTIO!i. 

AAOP MAGNITUDE OF THE PILOT OUTPUT SIGNAL ATTRIBUTED TO THE 

NUMERATOR OF THE PILOT TRANSFER FUNCTION. 

AAt CONSTANT FOR THE r:IRST ORDER NUMERATOR TERM OF THE AIRCRAFT TRAtlSFER FUNCTION. 

AA2 CONSTANT FOR THE SECOND ORDER NUMERATOR TERM OF THE AIRCRAFT TRANSFER FUNCTION. 

AA3 CONSTANT FOR THE THIRD ORDER NUMERATOR TERM OF THE AIRCRAFT TRANSFER FUNCTION. 

ABRTHT ALTITUDE ABOVE THE DECK MOTION MEAN AT WHICH THE PILOT 

IS TO ABORT 'CHASING' THE SHIP DECK AND RETURN TO THEASSIGNED HOVE~ ALTITUDE. 

B1 CONSTANT FOR THE FIRST ORDER DENOMINATOR TERt-l OF THE 

PILOT TRANSFER FUNCTION. 

82 CONSTANT FOR THE SECOND ORDER DENOMINATOR TERM or THE 

PILOT TRANSFER FUNCTION. SET'" 0.0 IN CURRENT PROGRAM 

BECAUSE NOT IN CURRENT USE. 

BAl CONSTANT FOR THE FIRST ORDER DENOMINATOR TERr-! OF THE 

AIRCRAFT TRANSFER- FUNCTION. 

BA1P~ MAGNITUDE OF THE PILOT OUTPUT SIGNAL ATTRIBUTED 'ro THE 

DENOIUNATOR TERM OF TilE PILOT TRANSFER FUNC'rION. 

B1\2 CONSTANT FOR THE SECOND ORDER DENOIUNATOR TEJU.l OF THE 

AIRCRAPT TRAtlSFER FUNCTION. 

BA3 CONSTANT FOR THE 'nIIRD ORDER DENOMINATOR TERM OF THE 

AIRCRAFT TRANSFER FUNCTION. 

A23 

~ 

I 

\ 

I 

~ )-

~ , , 

; , 

I 

~ 
'\', 

" .,/ .. ' 

BEG INSTRUC'r-ION ro PLOT3.FOR SUBROUTINE INDICATING 

BEGINNING POINT OF THE GRAPH X-l\XIS. 

BUF DIMENSIONAL DUMMY VARIABLE USED IN THE TIME SUBROUTINE. 

BUFF DIMENSIONAL DUMM'i VARIAJ)LE usm IN 'IHE DATE SUBROUTINE. 

Cl FIRST RECURSION CONSTANT USED IN A.I.L. CALCULATION OF 

PILOT MOD~ OUTPUT. 

C1A FIRST RECURSION CONSTANT USED IN A.I.L. CALCULATION OF 

AIRCRA~ MODEL. OUTPUT. 

C2 REMNANT. FROM EARLIER PROGRAM. NOT USED l",~ THIS VERSION. 

C2A SECOND ItECURSION CONSTANT usm IN A. I. L. CALCULATION. 

C2B 

SAME VARIABLE NAME IS USED FOR BOTH PILOT AND AIC MODELS. 

SECOND RECQRSION CONSTANT usm IN A.~.L. CALCULATION OF 

PILOT MODEL OUTPUT USED AS InITIAL CONDITION INPUT ro 

PILOT MODEL. 

C3A THJ.RD RECURSION CONSTANT USED IN A.I.L. CALCUL1\'rION. 

SAME VARIABLE NAME IS USED FOR BOTH PILOT AN6 AIC MODELS. 

C3B THIRD RECURSION CONSTANT USED IN A.I.L. CALCULATION OF 

PILOT MODEL OUTPUT USED AS INITIAL CONEHTION INPUT TO 

PILOT MODEL. 

C4A FOURTH RECURSION CONSTANT USEO IN A.I.L. CALCULATION. 

C4B 

SAME VARI~BLE NAME IS USED FOR BOTH PILOT AND A/C MODELS. 

FOURTH RECURSION CONSTANT USED IN A.I.L. CALCULATION OF 

PILOT MODEL OUTPUT USED AS INITIAL CONDITION INPUT TO 

PILOT MODEL. 

CSA,CSB,C6A,C6B,C7A,C7B,C8A,C8B ARE SIMIALAR TO THE ABOVE. 

CKPB REMNANT FROM EARGIER PROGRAM. NOT USED IN THIS V.ERSION. 

CTNH 'CHOP THROTTLE NOW HEIGHT'; HEIGHT ABOVE THE DECK AT 

\-lHICH THE PILOT IS ro QUICKLY REDUCE THRUST FOR LANDING. 

01 REHNANT FROM EARLIER PROGRAM. NOT USED IN THIS vERSION. 

02 REMNMT FROM EARLIER PROGRAH. NOT usrn IN TillS VERSION. 

03 AN OFTEN. US-ED COt1BINATION OF OTHER VA.RIABLES. 

A24 

\ 

" 
/ 

,.,...--., 

( 

'.~.' 

.-", ! ' 

------

DT\MP MAGNITUDE OF VERTICAL VELOCITY DAMPING THROUGH THRUST 

TERM. 

I:JH PILOT PRECIEVED E.."lROR IN HEIGHT (COMMANDED HEIGHT -

ACTUAL HEIGHT + NOISE). 

DHD PILOT PRECIEVED ERROR IN VERTICAL VELOCITY. 

DHDD PILOT PRECIEVED ERROR IN VERTICAL ACCELERATION. 

DllDTR TRUE ERROR IN VERTICAL vELOCIT7. 

DHTR TRUE ERROR IN HEIGHT. 

ELC ENGINE LAG TIME CONSTANT. 

EN RANDOM PHASE ANGLE GENERATED AND USED IN THE GENSM. FOR 

SUBROUTINE. 

ERR ERROR TERM (RANDOM NOISE WITH MGNlTUDE FROM -0.5 TO 

0.5) • 

ERROR 

FLAG 2 

FLT 

FLTl 

FPLS 

FPLSl 

ERROR TERH (RANDOM NOISE WITH MAGNITUDE FROM 0 TO 1). 

REMNAN'£ FRC»t EARLIER PROGRAM. NOT USED IN THIS VERSION. 

FLIGHT TIME (INITIAL. TIME MINUS TIME AT TOUCIIDONN). 

PREVIOUS FLIGHT TIME VALUE, HBLe FOR USE IN STATISTIC 

CALCULATIONS. 

'FLIGHT PATH LOGIC SLOT'; INTEGER USED TO SHOW SECTION OF 

FLIGHT PATH COMMAND LOGIC IN USE tmrn MONITERING 

SIMULATION. 

VARIABLE USED IN FINAL PRINTOUT TO INDICATE IF THE FLIGHT 

PATH COMMAND LOGIC HAD CYCLED THROUGH THE 'CHOP THROTTLE' 

SEQUENCE AT LEAST ONCE. 

GAP VARIABLE WHIcn DETERMItlES THE AMPLITUDE FOR A COSINE 

SMOOTHING FUNCTION IN THE FLIGHT PATH COMMAND LOGIC. 

GAPP VARIABLE NHICH DETERMINES THE AMPLITUDE FOR A COSINE 

SMOOTHING FUNCTION IN THE FLIGHT PATH COMMAND LOGIC. 

GO DUMMY VARIABLE USED TO TRANSFER THE VALUE OF GIIDOT 

BETWEEN VASCON.FOR ru.D VA1C.FOR. 

GHDOT GAIN USED IN CALUCULATING THE VERTICAL VELOCITY DAHPING THROUGH THRUST VALUE. 

A25 

\ 

" 
,-.... 

,. \ J.' , I, ( I' .. 

, ---, 

f., ~, 

U 

11 

n 

IT 

n 

r I 

I 

I 

I : j.: 

i 

J 

----------~------~'~ 

GPP VARIABLE WHICH DETERMINES THE AMPLITUDE FOR A COSINE 

SMOOTHING FUNCTION IN THE FLIGHT PATH COMMAND LOGIC. 

HLIM VARIABLE'WHICH CAN BE USED TO SET A LIMIT ON THE LOWER 

BOUNDS OF COMHANDED THRUST. 

HTA THE NUMERATOR PORTION OF THE PILOT MODEL OUTPUT USED AS 

THE INPUT CONSTANT FOR THE ZEROTH TERM IN THE NUMERATOR 

OF THE AIRCRAFT -TRANSFER FUNCTION AFTER ADDITION OF 

NOISE AND LIMITS. 

HTB THE DENOMINATOR PORTION OF THE PILOT MODEL OUTPUT USED 

HTPO 

AS THE INPUT CONSTAt."T FOR THE ZEROTH TERf.1 IN THE DENOMINATOR OF THE AIRCRAFT TRANSFER FUNCTION AFTER ADDITON OF 

NOISE AND LIIUTS. 

MAGNITUDE OF THE PILOT MODEL OUTPUT BEFORE THE ADDITION 

OF NOISE OR LIMITS. 

HTPOA FIRST PORTION OF EQUATION USED ~ CALCULATE HTPO. 

HTPOB SEC0ND PO.~TION OF EQUATION USED TO CALCULATE .HTPO. 

I DUMHY VARIABLE FOR DO STATEMENT. 

100 DUMMY VARIABLE FOR DO STA~EMENT. 

IFLAG1 0 LOGIC SWITCH usm TO CONTINUE 'CHOPPED THROTTLE' 

SEQUENCE IN THE FLIGHT, PATH COMMAND LOGIC AFTER ITS 

INITIALIZATION. 

IFLAG11 LOGIC SWITCH TURNED ON IN THE 'RUN FAST' SEQUIDlCE OF THE 

FLIGHT PATH COr,It'-IAND LOGIC AND USED TO DETERtlINE THE 

INITIAL AMPLITUDE FOR THE COSINE SMOOTHING FUNCTION I,N 

THE 'RUN' SEQUENCE OF THE FLIGHT PATH COMMAND LOGIC. 

IFLAG12 LOGIC SWITCH TURNED ON IN THE 'RUN' SEQUENCE OF THE 

FLIGHT PATH LOGIC AFTER INITIALIZA'rION OF TII.ffi CONSTANT 

FOR THE COSINE SM<YJTHING FUNCTION, TO PREVENT REINITIALIZATION O~ CONSECUTIVE PASSES. 

IFLAG13 LOGIC SWI'ICH TURNBO ON IN ,THE 'ABORT TO 

HOVER ALTI'l'UOE' SEQUENCE OF 'rHE 

FLIGHT PATH LOGIC AFTER mITrALIZATION OF TIHE CONSTANT 

FOR THE COSl~E SMOOTHING FUNCTION, TO PREVENT REINITIALIZATION ON CONS!':CliTIVE .PASSES. 

A26 
I FLAG. 3 LOGIC S\lITCH TURNED ON IN THE 'ABORT 'ID HOVER ALTITUDE' 

SEQUENCE OF THE FLIGHT PATH COMMAND LOGIC AFTER' INITIALIZATION OF A TIME CONSTANT FOR THE COSINE SMOOTHING FUNCTION TO PREVENT RE1NITIALIZATION ON CONSECUTIVE PASSES. 

IFLAG.5 DATA OUTPUT SWITCH, IF IFLAG5=1 THEN THE STATISTICAL 

INFORMATION IS CALCULATED AND OUTPUT. 

IFLAG.7 LOGIC SWITCH TUR~ED ON IN THE 'CHOP THROTT~E SEQUENCE' 

OF THE FLIGHT PltTH COMMAND LOGIC AND USED IN THE 'CHAS!::' 

SEQUENCE OF THE FLIGHT PATH COMMAND LOGIC TO DETERMINE 

THE INITIAL, AMPLITUDE FOR THE COSINE SMOOTHING FUNCTION. 

IFLAGB LOGIC SWITCH TURNED ON IN THE IN4.FOR SUBRQUTINE AFTER 

INITIAL COMMANDED LETDOWN FLIGHTPATH REACHES THE CONSTANT 

HOVER ALTITUDE VALUE, PREVENTS REINITIALIZATION OF LETDOWN 

SEQUENCE DURING. A RUll. 

IFLAG9 LOGIC' SWITCH TURnED ON IN THE 'CHOP THROTTLE' SEQUENCE 

OF THE FLIGHT PATH COMMAt."D LOGIC, WHICH CAN BE USED TO 

CHANGE THE THRUST LOWER LItUTS FROM THE NOtUNAL VALUE 

USED FOR THE REST OF FLIGHT PATH COMMAND SEQUENCES. 

INT 

ISEAS 

DUMMY VARIABLE USEQ IN DO LOOP. 

INTEGER VALUE OF SEA STATE BBING SIMULATED. IT IS 

PASSED TO VARIOUS SUBROUTINES AS A SIMULATION PARN1ETER 

AND FOR PARAMETER PRINTOUT FOR THE SIMULATION ON DAT~ 

OUTPUT. 

IT DUMMY VARIABLE USED IN VARIABLE DIHENSION STATEMENTS. 

ITt TIME INCREMENT USED' IN THE A.I.L. CALCULATIONS 

IT2 THROUGH ITB ARE TIME INCREHENTS USED IN THE A.I.L. CALCULATIONS. IT2=ITt**2/21, IT3=IT1**3/31, ETC. 

IT9 THROUGH IT12 ARE REMNANTS FROM EARLIER PROGRAMING. NOT USED 

IN THIS VERSION. 

ITOT TOTAL NUMBER OF ENTRIES USED IN DETERMIHG A HISTOGRAM OF 

STATISTICAL DATA. 

J INITIAL SEED FOR USE IN THE RANDOl1 NUMBER GENERATIllG 

SUBROUTINE. 

JAY DUMlol.Y VARIABLE FOR PASSING THE NUMBER OF GROUPS BEING 

RUN IN A GIVBl SIMULATION SESSION 'ID V~RIOUS SUBROUTINES. 

A27 

I 

/ 

i 

I 

! 

I 

I 

II 

f 

,I 

! -

I 

I 

"-

~, 
1'-" 

JRAN 

K 

KAY 

KDOO 

KMeAN 

KMN 

KP 

KPA 

KPB 

KPC 

L 

LIM 

.;. "., ' .. ~. 

INITIAL SEED FOR USE IN THE RANDOM NUMBER GENERATING 

SUBROUTINE. 

NUMBER OF RUNS/GROUP OF A SIMULATION SESSION. 

DUMMY VARIABLE, NUMBER OF RUNS/GROUP OF A SJMULATION 

SESSION. 

INCREMENTED VARIABLE USED AS A DIRENSION IN A D) LOOP. 

NUMBER OF ENTRIES IN A STATISTICAL MEAN CALCULATION. 

NUMBER OF ENTRIES IN A STATISTICAL ~ CALCULATION. 

NOMINAL PILOT GAIN. 

MODIFIED PILOT GAIN. 

MODIFIED PILOT GAIN. 

MODIFIED PILOT GAIN. 

INTEGER VALUE OF THE PILOTS PURE LAG TIME MULTIPLIED BY 

100. 

C~CULATED VARIABLE WHICH "'DDIFIES THE qT/W VALUE OUTPUT 

FROM THE PILOT MODEL TO KEEP IT IN PHYSICALLY REALIZABLE 

LI"'ITS. 

MIVRD VALUE OF THE MAXI~UM INSTANTANEOUS VELOCITY RELATIVE TO 

DECK ENCOUNTERED DURING A SIHULATION RUN. 

MTOUT MEAN THRUST/WEIGHT RATIO 0UTPUT DURING A SIMULATION 

RUN. 

POUT VALUE OF THE PILOT MODE~ OUTPUT, AFTER ADDITION OF NOISE 

AND LIMITS. 

PRECYC NUMBER OF CYCLES THE SUBROUTINE TRANO.FOR IS TO RUN 

THROUGH TO OBTAIN STEl.OY STATE OUTPUT BEFORE VALUES FOR 

NOISE AND TURBULENCE INPUTS ARE RETURNED TO THE MAIN 

PROGRI\M. 

RNP VALuE OF NOISE CALCULATE!) IN TRANO.FOR INSERTED I~4TO THE 

PILOT MODEL AS INTERoN/\L <,ILOT ·10I5E. 

RNQ VALUE OF NOlS E CALCULATED IN TFANO. FOR ADDEO 'It) THE 

POSITON ERROR (PILOT PREC:EPTI('~ NOISE). 

51 COt1MANDED FLIGHT PATH ALTITUDE. 

A:!B 

j 

, 

I ,. 

I 

! 

I 

, I, 

I 

i 

L 

I 

) 0 

~~ ! 

! 

, 

" 
SSGAP 

SX 

T 

Tl 

T2 

T4 

COMMANDED' FLIGHT PATH VELOCITY. 

COMMANDfn FLIGHT PATH ACCEL~TION. 

COMMANDED HOVER ALTlTUuE. 

COMMAIlDED HOVER VELOCITY. 

COMMANDED HOVER ACCELERATION. 

IN:::TIAL POINT FOR'A COSINE SMOOTHING FUNCTION IN THE 

'RUN' FAST' SEQUENCE OF THE FLIGHT PATH LOGIC. 

INITIAL POINT FOR A COSINE SMOOTHING FUNCTION :::~ THE 

'CHASE' SEQUENCE OF TIlE FLIGHT PATH LOGIC. 

nUTIAL POINT FOR A COSINE SMOOTHING FUNCTION IN THE 

'RUN' SEQUIDlCE OF THE FLIGHT PATH LOGIC. 

SHIP DECK POSITION.· 

RELATIVE DISTANCE BETWEEN SHIP DECK AND AIC DELAYED 

BY THE VALUE OF THE PURE PILOT LAG TUm. 

SHIP DECK VELOCITY. 

RELATIVE VELOCITY OF THE SHIP DECK AND AIC DELAYED 

BY THE VALUE OF THE PURE PILOT LAG TIME. 

SHIP DECK ACCELERATION. 

AMPLITUDE OF THE COSINE SMOOTHING FUNCTION IN THE 

'ABORT TO HOVER ALTITUDE' SEQUIDJCE OF THE FLIGHT PATH 

COMMAND LOGIC. 

INITIAL POSITION FOR THE START OF THE COSINE SMOOTHING 

FUNCTION IN THE 'ABORT TO HOVER ALTITUDE' SEQUENCE OF THE 

FLIGHT PATH COMMAND LOGIC. 

DUMMY VARIABLE IN CALL TO INPUT1.FOR, NOT USED IN THIS 

VERSION. 

TIME (SECONDS). 

110DIFIED VALUE OF T2, NOT USED IN THIS VERSION. 

REMl1ANT FROM EARLIER PROGRAM, NOT USED IN THIS VERSION. 

MODIFIED VALUE OF T~, NOT USED IN THIS VERSION. 

A29 
. ~'-' . I 

TC TIME INCREMENT FOR RUNNING THE SIMULATION. 

TE THROUGH TE4 TIME INITIALIZED AT THE BEGINING OF 1\ SET OF 

CONSECUTIVE Rm~S THROUGH ONE Of' THE FLIGHT PATH COMMAND 

LOGIC SEQUENCES FOR TIMING OF THE COSINE SMOOTHWG . 

FUNCTIONS. 

TF FINAL TIME FOR WHICH THE SIMULATION ABORTS A RUN IF 

TOUCHDOWN HAS NOT BEEN ACHEIVED. 

TFIN DUMMY VARIABLE USED IN GRAPHICS SUBROUTINE TO INDICATE 

X-AXIS MAXIMUM VALUE FOR GRAPH. 

TG PILOT LAG TIME CONSTANT. 

TGEE VALUE CALCULATED IN TRANO. FOR FOR INPUT AS TURBULENCE 

INTO THE AIC TRANSFER r~NCTION. 

TIl THE INITIAL TIME A RUN \lAS STARTED. AT AS OUTPUTTED FROM 

A RANDOM NUMBER GENRATION SEQU~~CE WITH BOUNDS FROM 0 

TO 3600 SECONDS. 

TL PILOT LEAD TIME CONSTANT • 

TOUT THRUST-TO-WEIGHT RATIO. 

TOUTMAX MAXIMUt1 THRUST-TO-WEIGHT RATIO COMMANDED BY PILOT (WITH 

AIC LIMITS) DURING A GIVEN RUN. 

TOUTMIN MINIMUM THRUST-'ro-WEIGHT RATIO COMMANDED BY PILOT' (WITH 

AIC LIMITS) DURING A GIVEN RUN. 

TSDA PART OF' THE SEQUENTIAL CALCULATION OF STANDARD DEVIATION 

FOR FLIGHT TIMES IN A GROUP OF RUNS. 

TW THE MAXIMUM THRUST-TO-WEIGHT RATIO ALLOWED FOR A GROUP 

OF RUNS. 

VSDA PART OF THE SEQUENTiAL CALCULATION OF STANDARD DEVIATION 

FOR TOUCH DO\'iN VELOCITIES IN A GROUP OF RUNS. 

VTD RELATIVE VELOCITY OF SHIP AND A/C AT roUCHDOWN. 

VTDl TOUCHDOWN VELOCITY FOR PREVIOUS RUN USED IN STATISTICAL 

CALCUL1'I.TIONS. 

VTZ DUMMY VARIABLE Pl\SSED TO SUBROUTINE TRANO. FOR. NOT USED IN 

THIS VERSION. 

11.30 

I \, 

i 

I 

\ 

I 

i 

~ 
" . 

..... /~ , 

I 

'i 

YO AN INITIAL CONDITION OF POSITION FOR THE,A.I.L. CALCULATION OF THE, DIFFERIENTIAL EQUATION DESCRIBInG THE PILOT 

TRANSFER FUNCTION. 

Yt 

YtA 

Yta 

Y2 

Y2A 

Y2B 

Y3 

'l3A 

Y3S 

'lAt 

'lAtA 

POSITION OUTPUT OF THE PILOT TRANSFER FUNCTION. 

FIRST PART OF POSITION CALCULATION FOR OUTPUT OF PILOT 

TRANSFER FUNC'rION. 

SECOND PART OF POSITION CALCULATION FOR OUTPUT OF PILOT 

TRANSFER FUNCTION. 

VELOCITY OUTPUT OF THE PILOT TRANSFER FUNCTION. 

FIRST PART OF VELOCITY CALCULATION FOR OUTPUT OF PILOT 

TRANSFER FUNCTION. 

SECOND PART OF VELOCITY CALCULATION FOR OUTPUT OF PILOT 

TRANSF,ER FUNCTION. 

ACCELERATION OUTPUT OF THE PILOT TRANSFER FUNCTION. 

FIRST PART OF ACCELERATION CALCULATION FOR OUTPUT OF 

PILOT TRANSFER FUNCTION • 

SECOND PART OF ACCELERATION CALCULATIml FOR OUTPUT OF 

PILOT TRANSFER FUNCTION. 

POSITION OUTPUT OF THE PILOT TRANSFER FUNCTION AS USE!) 

FOR INPUT TO AIC TRANSFER FUNCTION. 

FIRST PART OF POSITION CALCULATImJ FOR OUTPUT OF THE 

PILOT TRANSFER FUNCTION AS USED FOR INPUT TO AIC TRANSFER FUNCTION. 

'lAtB SECOND PART OF POSITION CALCULATION FOR OUTPUT OF TIlE 

PILOT TRANSFER FUNCTION AS USED FOR INPUT TO AIC TRANSFER FUNCTION. 

YA2 VELOCITY OUTPUT OF THE PILOT TRANSFER FUNCTION AS USED 

FOR InPUT TO, AIC TRANSFER FUNCTION. 

YA3 ACCELERATION OUTPUT OF THE PILOT TRANSFER FUNCTION AS 

USED FOR INPUT TO THE AIC TRANSFER FUNCTION. 

YAA' POSITIOtl OUTPUT OF TIm PILOT TRANSFER FUNCTION USED FOR 

MONITERING PILOT TRANSFER FUNCTIOtl OUTPUT. 

YAAl1\. FIRST PART OF POSITION OUTPUT OF PILOT TRANSFER FUNCTION 

CALCULATION. 

A31 

- -"''''- ,- ,.-, --:---------...::-..... 
I : 

IT 

r 

f· 

YAA1B SECOND PART O~ POSITION OUTPUT O~ PILOT TRANSFER 

FUNCTION CALCULAT!ON. 

YAA2 VELOCITY OUTPUT OF THE PILOT TRANSFER FUNCTION USED FOR 

MONITERING PILOT TRANSFER FUNCTION OUTPUT. 

YAA3 ACCELERATION OUTPUT OF THE PILOT TRANSFER FUNCTION USED 

FOR MONITERING PILOT TRANSFER FUNCTION OUTPUT. 

YDO INITIAL CONSTM1T FOR VELOCITY USED IN THE A.I.L. CALC~LATION OF THE DIFFERIENTIAL EQUATION REPERSENTWG THE 

PILOT TRANSFER FUNCTION. 

Z AIRCRAFT HEIGHT ABOVE MEAN DECK POSITION. 

ZAl FIRST PART O~ THE POSITON CALCULATION FOR AIC TRANSFE~ 

FUNCTION OUTPUT .. 

ZA2 SECOND PART OF THE POSITION CALCULATION FOR AIC TRANSFER 

FUNCTtON OUTPUT. 

ZBl FIRST PART OF nlE VELOCITY CALCULATION FOR AIC TRANSFER 

FUNCTION OUTPUT. 

ZB2 SECOND PART OF THE VELOCITY CALCULATION FOR AIC TRANSFER 

FUNCTION OUTPUT. 

ZCl FIRST PART OF THE ACCELERATION CALCULATION FOR AIC TRANSFER FUNCTION OUTPUT. 

zci SECOND PART O~ THE ACCELERATION CALCULATION FOR Ale 

TRANSFER FUNCTION OUTPUT. 

ZO AIC VELOCITY OUTPUT. 

zoo AIC ~CCELERATION OUTPUT. 

ZW VERTICAL VELOCITY DAHPING DUE TO AIRFRAHE TIME CONSTAN'r. 

ZWDTDT NUMERATOR CONSTANT FOR THE AIC TRANSFER FUNCTION. 

A32 
r [ 

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[ 

[ 

I~ 

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. ', 

I: 

LIST OF VARIABLES FOR IN4.FOR 

NAME DESCRIPTION 

A THE INITIAL TIME FOR THE • A' SEQUENCE OF THE HOVER ALTITUDE COMMAND LOGIC. 

AHPA FINAL HOVER COMMAND HEIGHT. ALSO THE AMPLITUDE FOR THE 

• A' SEQUNCE OF THE HOVER COMMAlID HEIGHT LOGIC (THE 

COSINE FUNCTION OF ' A' SEQUENCE IS NOT USED IN THIS 

VERSION) • 

AMPB THE AMPLITUDE FOR THE COSINE FUNCTION OF THE • B' SEQUENCE" 

OF THE HOVER HEIGHT LOGIC. AMPA+AMPB GIVES THE INITIAL 

STARTING HEIGHT ABOVE THE" SHIP DECK MEAN. IN THIS VERSION. 

B THE INITIAL TIME FOR THE • B' SEQUENCE OF THE HOVER ALTITUDE COMMAND LOGIC. 

IFLAGS LOGIC SWlnH TURNED ON AFTER THE INI'rIAL COMMANDED LETDOWN 

FLIGHTPATH (' B' SEQUENCE) REACHES TrlE CONSTANT HOVER ALTITUDE VALUE (AHPA) TO PREVEUT REINITIALIZATION OF THE 'B' 

SEQUENCE DURING A GIVEN RUN. THIS VARIABL~ CORRESPONDS 

TO IFLAG8 IN TIlE IN4.FOR CALL STATEMENT FROM VPAIC.fOR. 

ISFLAG1 LOGIC ~WITCH TURNED ON AFTER THE TIME INITIALIZATION IN 

THE 'A' SEQUENCE TO PREVENT REINITIALIZATION OF THE TIME 

DURING A GIVEN RUN. 

Sl COMMANDED HOVER FLIGHT PATH ALTITUDE. 

S2 

S3 

SX 

COMHANDED HOVER FLIGHT PATH VELOCITY. 

COMMANDED HOVER FLIGHT PATH ACCELERATION. 

DUMltY VARIABLE usrn IN CALL STATEMENT TO IN4. FOR, NOT 

USED III TIUS VERSION. 

T TIME (SECONDS). 

TAU TUm HINUS THE TIME OF INITIALIZATION OF EITHER THE 'A I OR 

'8' SEQUENCE. 

TI THE INITIAL TIME A RUN \'lAS STARTED AT, AS OUTPUTTED FROM 

A RANDOH NUMBER GENERATlml SEQUENCE WITH BOUNDS FROI-t 0 

TO 3600 SECONDS. 

A33 
,,'. 

r\, \ I; 

, 1 

1\ 1 • 

TIMA VARIABLE USED IN 'mE INITIALIZING. THE TIME FOR THE 'A' 

SEQUEtICE. 

WNS FREQUENCY OF THE COSINE FUNCTION IN THE 'A' SEQUENCE. 

WNSl FREQUEtlcr OF THE COSlNE FUNCTION IN THE 'B' SEQUENCS. 

LIST OF VARIABLES FOR TRANO.FOR 

NAME DESCRIPTION 

FLAG REMNANT FROM EARLIER PROGRAM, NOT USED IN THIS VERSION. 

I DUMMY VARIABLE USED IN 00 LOOPS. 

IB IFI;'{ED VALUE OF XB. 

IS IFIXED VALUE OF S. 

H THE NONINTEGER PART OF THE TIME (IN SECONDS) OF THn 

SYSTEM cLOCK. 

PRECYC NUMBER OF CYCLES THE TRANO. FOR SUBROUTWE IS TO RUN 

THROUGH TO OBTAIN STEADY STATE OUTPUT BEFORE VALUES FOR 

NOISE. AND WRBULENCE INPUTS ARE RETURNED TO. THE !-!AIN 

PROGRAM. 

RNO OUTPUT OF TIlE" FIRST ORDER FILTER USED AS INTERNAL PILOT 

NOISE. 

RNOO FIRST DERIVATIVE OF RNO. 

RNP RNO MODIFIEC BY A GAIN. 

RNQ OUTPUT OF THE FIRST ORDER FILTER USED AS PILOT PRECEPTION ERROR. 

RNQD FIRST DERIVATIVE OF RNQ. 

RT SYSTEM TIME MINUS THE nON INTEGER PART OF THE SYSTEM 

TIM.!':, IN SECONDS. 

S INDICATES IF. XM IS EVEN OR ODD. 

A34 
/ 

SIGMA· 

SIGMAN 

..... - .' 

THE SIGMA VALUE FOR THE FILTER USED IN GENERATING RANDOM 

TURBULENCE. 

THE SIGMA VALUE FOR THE FILTER USED IN GENERATING THE 

ItiTERNAL PILOT NOISE. 

SIGMAQ THE SIGMA VALUE FOR THE FILTER USED IN GENERATING THE 

PILOT PRECEPTION NOISE. 

TC TIlE PROGRAM TIMf: INCREMENT. 

TGEE OUTPUT OF THl:: r·t<IST ORDER FILTER usm TO GENERATE: 

RANDOM TURP,uLENCE. 

'I'M SYSTEM 'tIME IN SECONDS. 

VTZ SAME AS 'roEE. 

WN FREQUENCY TERM FOR THE 'tuRBULENCE SHAPING FILTER. 

WNN FREQUENCY TERl-i FOR THE INTERNAU PILOT NOISE SHAPING 

FILTER. 

WNQ FREQUENCY TERM FOR THE PILOT PERCEPTION ERROR SHAPING. 

FILTER.' 

XB MODIFIED VALUE OF RT. 

XM FLOATm VALUE OF M. 

XS FLOATEV VALUE OF IS. 

Y VALUE. OF 12 RANDOM NUMBERS ADDED TOGETHER IN THE PROCESS 

OF CREATING A GAUSSIAN DISTRIBUTION FROM A WHITE NOISE 

SOUR~E. 

YA A RANDOM NUHBER GENERATED FROM A WHITE NOISE SOURCE. 

YB A RANDOM NUMBER GENERATED FROM A GAUSSIAN OIS~RIBUTION 

SOURCE. 

A35 

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....... 

. I 

\ 

.. 
'. 

.~, 

:.! 

/ 

.r 

, 

\ 

" 

VARIABLES FOR GEN3M.FOR 

NAME DESCRIPTION 

EN RANDOM PHASE ANGLE. 

Hl THROUGH H6 ARE THE CON.TRIBUTIONS ro THE SHIP HEAVE IN FEET, 

FROM EACH OF THE SIX SlUE CONPONENTS ro THE SHIP 

HEAVE MOTION APPROXIMATION. 

HDl THROUGH H06 ARE THE CONTRIBUT!ONS ro THE SHIP HEAVE VELOCITY 

IN FEET/SEC, FROM EACH OF THE SIX SINE CONPONENTg ro 

THE SHIP HEAVE MOTION APPROXIMATION., 

HDOl THROUGH HODS ARE THE CONTRIBUTIONS ro TH~ SHIP HEAVE ACCELERATIOfl IN FEET/SEC**2, FROM EACH OF TH,E SIX SINE CONPONENTS OF THE SIX SINE CONPONENTS ro THE SHIP HEAVE 

MOTION APPROXIMATIC:I. 

IB IFIXED VALUE OF XB. 

M IFIXED VALUE OF S. 

Pl THROUGH P6 ARE THE CONTRIBUTIONS ro,THE SHIP PITCH POSITION 

IN RADIANS, FROM FACH OF THE SIX SINE CONPONENTS TO 

SHIP PITCH MOTION. APPROXII-!ATION. 

PD1 THROUGH PD6 ARE THE'CONTRIBUTIONS ro THE SHIP PITCH VELOCITY 

IN RADIANS/S~C, FROM EACH OF THE SIX SINE CONPONIDJTS. ro 

SHIP PITCH MOTION APPROXIMATION. 

POD1 THROUGH PDD6 ARE THE CONTRIBUTIONS ro THE SHIP PITCH ACCELERATIOi~ IN RADIANS/SBC**2, FROH EACH OF THE SIX SINE 

CONPONENTS ro SHIP PITClI ~IOTION APPROXIMATION. 

RT SYST~1 TIME MINUS THE NONINTEGER PART OF THE SYSTEM 

THIE IN SECONDS. 

S INDICATES IF XM· IS WEN OR ODD. 

Sl SHIP DECK POSITION RELATIVE ro n1E MEAN DECK POSITION. 

S2 SHIP DECK VELOCITY. 

S2A SUM OF THE SHIP HEAV·E VELOCITY CONTRIBUTIONS TO 'TIlE 

SHIP HEAVE HOTtON. 

A36 

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J 

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• 

1 

L 

. , -"" " 

'. 
, 

, 

~J I.. 

rJ 

II 

u 

11 

II 

E 

I 

i 

I 

S2B SUM OF THE SHIP PITCH VELOCITY CONTRIBUTIONS, MULTI-. 

PLIED BY THE DISTANCE FROM THE SHIP C.G. TO THE SHIP 

LANDING PAD BULLSEYE TO OBTAIN THE CONTRIBUTION OF THE 

SHIP PITCHING MOTION TO HEAVE AT THE LANDING PAD BULLSEYE. 

S3 SHIP DECK ACCELERATION. 

S3A SUM OF THE SHIP HEAVE ACCELERATION CO~TRIBUTIONS TO TIlE 

SHIP HEAVE ACCELERATION. 

S3B SUM OF THE SHIP PITCH ACCELERATION· CONTRIBUTIONS, MULTIPLIED BY THE DISTANCE FR9M THE SHIP C.G. ~O THE SHIP 

LANDING PAD BULLSEYE TO OBTAIN THE CONTRIBUTION OF TIlE 

SHIP PITCHING HOTION TO HEAVE AT THE LANDING PAD BULLSEYE. 

T 

T1 

TM 

TIME (SECONDS). 

THE INITIAL TIME A RUN WAS STARTED AT, AS OU1~UTTED 

FROM A RANDOM NUMBER GENERATION SEQUENCE WITH BOUNDS 

FROM 0 TO 3600 SECONDS. 

SYSTEM 'Tlt{E IN SECONDS. 

XB MODIFIED VALUE OF RT. 

XM FLOATED VALUE OF M. 

XS FLOATED VALUE OF IS. 

Y RANDOM NmmE"~ \HTIl MAGNITUDE OF -0.5 'ro 0.5. 

A37 

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DllIDISltN $1.aT(50). TJ.lI:1R(50) ,IS(50) ,W1S(SO) 

DIMENSltN \ffi(5QO) ,~(500), ntl(500) , lSD(500) ,TI~(500) 

~(;.L K!' .1T1.1T2.ITJ.1T4 • ITS. m,m ,IT9,IT9,1TI0, !Til 

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(;.LL Tt;iiSLlQ.t((TG££, TC.IJTZ ,PPLC'I'C,f:1P ,m'l) 

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CALL INPUTI(SA1.SAZ,SAl,I,III,IFtAGS,SX) 

CALL IHftlT2(SS1,SS2.SS3, t,III,Di,~.A!.PIIIZ,ATH.PHIO) 

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SSlZ=.SSl-Z 

SSZZMS2-ZO 

IF (lFlMlO .EO. 1 .t#). T£4 .LT. 1.5) C-O TO 50 

IF (552 .Gr. 5.5) GO TO 3S 

If (SSIZ .GT. -9.0 .W'"!. SS2 .Gr. 4.5) GO TO 43 

IF (SSlZ .LT. -3.0) (j() TO 45 ,.. .'\..- IF (Z .IT. ABl!TH) GO TO 45 

If (SSlZ .Gr. CTNH) CO TO 50 . - . I -

IF (lFLAG3 .EO. 0) TIOl 

TEl:O.O 

IrLAG3=1 

I FlAG7=O 

IFLAG9=O 

IrLAGIO=O 

IFlAGll-O 

IFtAGl2=O 

lF1.AGIJ:O 

~:;P-Sl 

CTN!i=-1.0 

Ern IF 

Sl-SSIH (5{,P-SSI )HSGP-SSI )*COS(6.2S32tTEl) )12.0 

52=552 

S3=SS3 

KP=l:PB 

If mPB .EO. 1) KP=tFA 

TEl=T£ltT~ 

If (TEl .Gr. 0.5) TEl=O.5 

FPLS'3 

GO TO 65 

35 If (IFtl\Gll .ro. 0) THEH 

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0182 

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0184 

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0189 

0190 

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0198 

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0200 

0201 

0202 

0203 

0204 

0205 

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0207 

0208 

0209 

0210 

0211 

0212 

0213 

0214 

0215 

0216 

0217 

0218 

0219 

0220 45 

0221 

0222 

0223 

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02:25 

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21-JiwH9S5 12:0~:3~ \%\X-II Fomn! V3 •. H2 Pi9' 

m:o.o 

IFlA63:O 

lflAG7zjJ 

I~ 

1f1AG10:O 

IflI,GlI·I 

IflAGl2:0 

IflA613:0 

SWP=SI 

WPs6.0 

CTIfla-l.0 

00 IF' 

51 :SfAI'P+(WP-fIJ'PtCOS(' • 2832*TE3) )/2.0 

S2>SS2 

SloSSl 

r,pom: 

T£3:TE3nc 

IF' (TE3 .IIT. 0.5) TEM.S 

fPlS=l 

60 TO ~ 

IF' (If1AG12 .EO. 0) TI£H 

TE:O.O 

I~ 

1FlJl67=O 

1f'lJ.G3:0 

IF!AG10=O 

IF' urtAGll .EO. 1) M:N 

SGl'PoSl 

GPP·-3.0 

US( IF (IflJ.GlI .EO. 0) TIel 

S{;PP'51 

GPP=3.0 

!}{) IF 

I FlAGlI =0 

IFlAG1Z.1 

1flJ.G13=O 

ClnF-1.0 

CO IF 

Sl·SU1't(GPP-GPPtCOS( 6.2832* TEl )12. 0 

S2=SS2 

SloSS3 

Xl'=KFS 

TE·TEtTC 

IF err .IIT. 0.5) TE=O.5 

fPlS-:2 

GO TO 65 

IF (Jrtf.(j13 .[0. 0) TIflI 

TE2=O.0 

SS(;;P:Sl 

S!';W':SAI-SSW 

I FlAG3:0, 

IF1AG9aO 

IFlAG10·O 

IFlAGlI=O 

JrlAGlN 

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

19-J~-1~5 21:06:48 fSOQ:[ST(l.UlS.SH[PSTLFF)VPAlC.FOR;17 

A65 

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0277 

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47 

50 

65 

70 

t. . ... " . 

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21-JiII-19a5 12:09:35 

U-JI!I-lm 21:0&:48 

1F1Ali13:1 

ctMto-l.0 

00 IF 

If (lFlAG7 .EO. 1) GO TO 46 

S1·5smPt(S5m'-S~('.W2*TE2»/2.0 

GO TO 47 

510W 

S2=SA2 

S3--W 

If (ltO.S .IT. 51) TftJI 

KP=KPa 

ElSE Ir mO.5 .6£. 51) TllElf 

KPzKI'II 

EN> If 

ru=mtTc 

If (TE2 .GT. D.5) Tlal 

T[2=O.5 

I FlAG7'1 

END IF 

fPl.S=O 

GO TD. 65 

IF (lFU'.G10 .EO. 0) TllDI 

IC.MM 

fPl.S-~ 

IFlI:G3--t 

1nAS7'1 

1rtAG9:1 

IFlAGlO=1 

I F1.t.G11 =0 

IF1.I.S12=O 

I FlAG13=0 

ClIfla-l.7 

T£4=O.O 

GAP=S1 

00 IF 

!l1r.=-I.O 

IF (m .LT. O.2S) TOOl 

S1=wm.ttCOS(12.~tTE4)-9.0)12.0 

FPlS=6 

ElSE If (m .6£. 0.25 .ltt). TE4 .IT. 1,2S) mH 

51 :fI1I'-9. 0 

FPlS:7 

ElSE IF (T£4 .f.[. l.z:i) mN 

51=WH9.ttCOS(12.56S4t(T[4-1»-9.0lI2.0 

FPl.s:a 

00 IF 

TE4:mut. 

IF ((SSI-Z) .6£. 0.0) GO TD 1~ 

00 70 l:l.2.-1 

$lS(I) =$lS( 1-1) 

$lV( I j.=$lV( 1'1) 

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manu: 

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0300 

0301 

0302 

0303 

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0305 

0306 

0307 

0309 

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0310 

0311 

0312 

0313 

0314 

0315 

0)'.6 

0317 

0319 

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0320 C 

0321 

0322 

0323 

0324 

0325 

032' 

0327 C 

0329 

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0330 

0331 

0332 

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0334 

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21-J~'19:l5 12:09:35 VAX,-Il fORT~ v3.5-62 PI9f 6 

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SlS(1l:S1 

Slvm·S2 

SlAm=S3 

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A6=dH 

AI-YO 

BlzlITG 

f2=O.O 

a=A1 

C2Ss-BltCI 

C3S'-B1tt2B-B2*C2 

C4!!=-B1*C3B-B2*C39 

C5B=-B1tC~B-B2*C4B 

C6Bs-sttC5B' J2*C5B 

C;'B:-BltC6B-B2*C68 

C8B=-B1tClB-82*ClB 

C2A·1.0 

o\=-BI tC2A' 

C~'-SI *Cl'I 

~-Bl*C~Hl2*C"A 

C6A'-BltCi-B2kC5\ 

C7A·-ette6.~·B2*C6A 

CMa-BltC7A-B2*C7A 

YlAsa tC2S* IT1 t03* I T2+C4Bt I T3t~Bt IT 4fC6BtI T5 

YlB=C7E* JTOC8!l* IT7 

Y2A:C23+C3Bt I 11 tc4!:* I T2+C5£i IT3tC6e* IT 4 

Y2!l=ClBt IT5tC8Bt m 

Y~C38tC4Bt IT! tC5Si IT2fC6Bt I T3 

Y38=C7Bt IT msBt I T5 

Y4A=C48tC5BtITlmB*1T2+C7BtIT3+C8BtIT4 

YAIA=C2A*11ItC?AAIT2+CWIT3tC5\tIT4+C6AtIT5fC7A*IT' 

YAIB=CM*1T7 

YAl'YAlMYAlB 

YA2=C2A+C?A* I 11 tCv.* I T2tC~ I mC6A* IT 4+ C7A* I T5fCM* IT' 

YA3=O'ttCv.t I T1 +C:;';l I me,Al I 13+C7A* I T 4tt8A* I 15 

YA4=eWC5\tlT1+C6At ITZ+C7MIT3+CeA*IT4 

C2A=14 

c:r.--BltC2A 

C4A'-B1t~ 

CSl.-Bl tev.-e2*C~ 

C6A'-BI *C5A-B2*C~ 

C7A'-Bl*C6A-Il2*CtA 

C&1'-Bl*C7A-B2*e7A 

YMlA=C2Atl 11 tCWIT2WA* JT3tC5\* I mCtA' I T5tC7A* lT6 

YMI B=C&'t* JT7 

YMI'YMlAtYMlB 

YM2~Cl'\t IT1 tt~* IT2tCW I mCtA* I T 4+C7A*1 T5fCtA*IT6 

YM3=O'ttCV.t lT1 tC~t I 12fCf.A* IT3+C7i\t I 1 4tCSAtI 15 

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j. 

". 

. I:' , 
, : 

--

,/-

i 

! 

.:; 

----

r 

I 

~ 

I /' 

II - Ii 

." " . ..... 

0343 

0344 

0345 

om 

I' I • 

0347 C 

0348 

034' 

~3S0 

O~ 

0352 

0353 C 

O~.A 

0355 C 

03$ C 

0357 C 

0~8 

0359 C 

0360 

0361 C 

0362 

0:1£3 

03U 

0365 

0366 

0367 

0368 

036' 

0370 

0371 C 

0372 C 

0373 

0374 74 

0315 

0376 

03n 15 

0378 

0379 

OJCil 

0381 C 

0382 

0:?83 80 

0334 

0385 

om 

0337 

03S8 

0383 

om 

om 

0392 

om 

om 

om 

om 

0331 

om C 

om C 

i ; 

/I' -/ I 

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, .,.. i",·' 

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21-JM-m512:0':jS ""'-11 rcmm V3.5-62 P31f 

Yl'YIAtYlB 

Y2aY2MY23 

YJoY3A+Y3S 

Y4'Y4AtYA4 

Dl'UlTl-11T2) 

D2a1l(Tlt-T2) 

D~( Kl'tTVTGJ 

HtS'YA2t(l/TlltYAl 

HiBaCYlUfJ)ltYA2-02tYAl) 

HtA:( Y2t$lS(L) tYAZI t( l/Tl)t (Y1 t$lS( L )iYAl) 

INM-lm 21:06:48 fSOO:[STEVENS.SHIPSTlJFF)Vl'AIC.rCR;17 

HIA: (Ylt$lS( l I tYAl) tOI *( Y2t$lS( L).tYAZ) -D2t( Yl t$lS( L I iYAl) 

Htf'0!3:-02t(Y1 tYMI) 

Ht~(Y3tYAA3)tOli(Y2tYM2) 

HtPOO=(lIR)t(Ylt($lS(L)-ZLCL)ltYAll 

HtPOOo-02t1Yl t( $lS( Ll-ZLCLl ItYAl l 

HtPU'\'Y2t(SLS(Ll-ZL(LlliYA2 

HtP()\:(Y3+( nS(L) -ZL(Ll I LYAlltOl t(.Y2t ($lS( L )-Zl( l) ItY.~) 

HtPO=03t(Htf'lil+HtPOB-) 

YO:YHYAAl t9.NP 

YOO·Y2tYAA2 

L111=1.0 

IF (HtPO .EO. 0.01 HlPO·D.DDOCl 

IF (J FLAG3 .EO. 1) GO Te 74 

11L1M·-((T1HI*1.51 

IF (HLIH .GT. -0.05) htll1=-D.05 

11.:1/1=-1.0 

IF (HtPO .GT. ~lIHI GIl TO 15 

1I11=!'l.IIVHrPO 

GO TO ~ 

IF. (HtPO .LT. (TIl-I)) GO TO £4 

lIH'(n!-l )/HtPO 

UUHHUU AlC T.F.F. tuuuuu 

MO'( (Zl-IOTDTt (ITG£E HW1P)ILIHI )IUC)tLlHtOltHtA 

AAl'ZtZWELC t ZOt (1/[l.C' 1lI1 t ZOO 

M2:Zi(VELCflllltZO 

M3'Z 

~1'( (ZlIOTOT t (CTG££ftw1PIILIH) )IELC) tLllt.tOltHtB 

~2·ZWE!.C 

&\3:Zllf lIELC 

MOP'lllt*03tKtA 

&\IP'lllt*OltHtB 

PO'JT:MOPt&'ll?t( -Zl(L)) 

TOIJT=PO!ITH 

IITOUT· ((Y.HEm-ll tHTM HO'JT )/Y.'IEm 

IF CTO!iT .GT. TO'Jn'ilX) TOOTmX.TOUT 

IF ITM .LT. TC'JlHINI TO!ITHIN:TO'JT 

IF ITO!iT .LE. 1.21 .fllD. TO!iT .61. 1.101) Tli£N 

GO TO as 

A68 

./ 

/ 

-j' 

I 

I 

u: 

{ 

I 

'; 

I 

, J . j 

. 

, 

I 

I 

I 

I 

I 

; 

r 
: 

! 

i 

! 

r, 

/, 

i 

t 

I 

,j 

j 

l 

..... 

l 

, 

., 

l 

0400 

am 

0402 

oe03 

0404 

0405 

0406 

0407 

0408 CB5 

0403 C 

0410 C 

0411 C 

0412 C 

0413 C 

om c 

0415 C 

C416 C 

0417 C 

0419 C 

om c 

042' C 

0421 C 

C422 C 

0423 C 

C424 C 

0425 C 

0426 C 

0427 C 

0429 C 

om C 

0430 sa 

0431 

0432 

0433 

0434 

0435 

0431i 

0437 

0433 

om 

0440 

0441 

0412 

0443 

0444 

0445 

044& 

0447 

0448 

044~ 

0450 

0451 

0452 95 

0453 

0454 

om 

0456 

/ . I : 1 ! 

--~-------" .. --...... 

IF (TOUT .tr. 1.101 .mo. TOUT .Gr. 1.001) THEN 

GO TO" 

21-J.H985 12:09:35 w.lC-U FOI!T~ V3.5-6Z Pi9' 8 

19-Jan-l985 21:06:4a rSOO:[STEVOIS.SHIPSTtJrF)VPAIC.rOl!;17 

ElSE IF (TQ,,7 .LE. 1.001 .ttlJ. TOUT .Gr. 0.9011 TI(N 

GO TO 95 

ElSE IF (TM .LE. 0.9\11 .ttlJ. TOUT .Gr. o.sOOI TH£N 

m TO 100 

00 Ir 

GO TO 105 

IF erM .LE. 1.201 .fW. TOUT .Gr. 1.ISl) THDf 

IS(1)'IS(1ltl 

ElSE IF (TM .tE., I.U! .fW. TOUT .GT. 1.181) TtIfH 

15(2)-15(21+1 

ru£ IF (TM .LE. 1.181 .rH). TOUT .GT. 1.1m THDf 

IS(3)=ISC3Hl' 

ElSE IF HM .LE. 1.171 .r«>. TM .GT. 1.161) TH£N 

IS(C)=IS(4lH 

ELSE IF (TM .LE. 1.161 .fH). TM .GT. 1.151) T~ 

IS(5)'15(SIH 

ElSE IF (TOUT .LE. 1.151 .tn). TO'JT .GT. 1.141) TH£N 

15(6):15(6)+1 

ELSE Ir (TroT .lE. 1.141 .A·~. TOUT .GT. 1.131) TflH 

IS(7)·IS(7)tl. 

ElSE IF' (TOOT .LE. 1.131 .rH). TOUT .GT. 1.121) TH£N IS(B)=IS(~It1 

ElSE IF (TOUT .LE. 1.121 .m;). TM .GT. 1.111) TH!N 

IS(S)=IS(9)tl 

ElSE IF (TM .LE. 1.111 .tllI. TM .GT. 1.IOU TH£N 

15(1O)=15UOHI 

En>lr 

GO TO 105' 

IF (TM .LE. 1.101 .rH). TOUT .GT. 1.091) TH!l/ 

1S(11)=1 S(11)tl 

ElSE IF (TM .LE. 1.091 .ftD. TOUT .GT. 1.081~ TH£N 

IS(12)=1 SC12lH 

ElSE IF' (TOUT .LE. I.OBI .1111). T:1JT .GT. 1.071) TH£N 

IS(J31=ISC13lt1 

ElSE IF' (TOUT .LE. I.on .fW. T'JUT .Gr. 1.061) TIlEn 

IS(J4)=IS(J4)t1 

ElSE, IF (Tour .LE. 1.061 .fW. Tour .GT. 1.051) THElI 

IS(J51=IS(J51tl . 

ElSE IF (TOUT, .LE. 1.051 .ft{). TM .GT. 1.041) TflH 

IS(J6)'IS(16)t1 

ElSE IF (TOUT .LE. 1.041 .ftlO. TM .GT. 1.031) Th!N 

IS(7)=ISCl7)t1 

ElSE IF (TOUT .LE. 1.031 .ft~. TO'JT .GT. 1.021] THDf 

IS(J9)=ISCl9)+1 

ElSE IF (TOUT .LE. 1.021 .(t,'O. TM .GT. 1.011) THEN 

IS(19)=lsm)+1 

ElSE IF (Tour .LE. 1.011 •• UJ. TM .GT. 1.001) THEN 

IS(20):ISC2O)tl 

En> IF 

GO TO 105 

IF (TOUT .LE. 1.001 .A'i!>. TOUT .GT.' 0.991) Tff}{ 

IS(21)cISC21)tl 

ElSE IF' (Tcx,'T .LE. 0.991 .tW, TOUT .GT. 0.991) THOI 

IS(221'IS(22)t1 

ELSE IF (TM .LE. O.~I .A'D. TM .GT. 0.971) Th1}I 

A69 

\ 

........A.'1 ~ '~ " 

'I I, 

.' ~ 

I 

• I k..J 

.", 

I 
,// 

I 

! 

~"!n' 

0457 

04~ 

0453 

0460 

0461 

0462 

0403 

0464 

0465 

046& 

0467 

0468 

0409 

0470 

0471 

0472 

Gtn 

0474 

0415 

om 

0477 

0478 

om 

04SO 

0491 

0482 

0483 

048.4 

0485 

0486 

0487 

0488 

0489 

0490 

0491 

om 

0433 

0494 

0495 

04% 

0497 

0498 

om 

0500 

0501 

0502 

0503 

0504 

0505 

0506 

0507 

0:08 

050' 

0510 

0521 

0512 

0513 

100 

IDS 

21-JiII-19S5 12:09:35 UAX-U fORTIWI V3.5-62 Pi9r 9 

19-JiiIl-19~ 21 :06:48 FSOO: [STOOIS.SlII PSTI.H]VPAIC.FOR; 17 

15(23'-15(23'+1 

ElSE IF' (TOUT .LE. 0.971 .tW. TOUT .GT. 0.961) THEN 

IS(24'=IS(24)+1 

ElSE If (TOUT .LE. 0.961 .mo. TOUT .GT. 0.951) TlIH. 

15(25)-15(2Sltl 

ElSE IF' (TOUT .LE. 0.951 .tn). TOIfT .GT. 0.941) TIm 

/5(26)=15(26)+1 

ElSE. IF (TWT .LE. 0.941 .tK!. TOUT .GT. 0.931) TID! 

IS(27)=15(27lt1 

ElSE IF (TOUT .LE. 0.931.i'tID. TOUT .GT. 0.921) TI£N 

15(23':./5(29)+1 

ElSE If (TOUT .LE. 0.921 .ftl!l. TOOT .GT. 0.911) TIIm 

15(29'=15(29)t·1 

ElSE IF (TIXIT .LE. o.m .m>. TOIfT .GT. 0.901) OOt 

15(3O'=15(30'tl 

00 If 

GO TO 105 

If (TOIIT .LE. 0.901 .fl!). To'.[r .GT. 0.991) THIN 

IS(31'=15(31'+1 

ElSE IF (TOUT .LE. 0.991 .tn). TOIfT .GT. O.BSI) TlI(N 

IS(32'=15(32)+1 

ELSE IF' (TOUT .LE. O.SSI .tn). TOlIT .GT. 0.871) 001 

15(33)=IS(33'tt 

ELSE IF' (TOIIT .LE. 0.871 .1tI!l. TOIIT .GT. 0.861) 001 

15(34'=IS(341+1 

ElSE. IF' (TIXIT .LE. 0.861 .mo. TOUT .GT. 0.851) TID! 

IS(35)'15(35'tl 

ElSE IF' (TOUT 'LE' 0.&51 .f.,{) •. TOLi .GT. 0.941) THIN 

IS( 36,=ISC;6ltl 

ELSE IF' (TOUT .LE. 0.8~1 .tn)". TOIfT .GT. 0.831) TI§N 

15(37)=15(37)H 

ElSE IF' (TIXIT .LE. 0.831 .tW. TM .Gr. o.e21) THEN 

IS(38)=15(39'tl 

ElSE. IF' (TOIIT .LE. 0.821 .tW. TOUT .Gr. 0.811) THEN 

IS(39'=15(39)tl 

ElSE IF' (TOIIT .LE. 0.811 .rtf). TOUT .GT. O.SOI) TIlm 

IS(40)=15(40'+1 

00 IF' 

ClA~ 

C2A=M2-SA3*ClA 

oww -E\42tCIA-~3*Q\ 

C4AWoIl-ft.l tClA-a '2tC2A-8i\3tC3~ 

c:A'-SA! *C2A-I:A2*C3A-SAltC4.<I 

CSA=-SAltC~-El42tC4A-rAltC5.\ 

C7A'-&'<! t C4A-E\42tC5A-SAltCSA 

C2A=-£l<U*C5.\-E\42tCSA..r.>\ltC7., 

ZAI :Clft! ClAt I TI tC:l.<t I mcw I mCSAt I T mtAt I T5 

W'C7At IT6tC(lA* IT7 

ZSI :C2il+C~ ITl fC!,i>t I T2tC~ I mcw IT 4 

ZB2=C7At lT5tCrAtJT6 

ZCI :C3AtCW ITI tC5At IT2tCW I T3 

ZC2=C7At IT 4tCPn"t IT5 

ZaZA1+W 

ZO=ZSHZB2 

A70 

I 

1 • " 

\, .. :,' 

, 

l 

. ~ 

J 

; 

I 

1 

l' . 
( 

WIS1P 

0514 

0515 

0516 

0517 

0518 

051' 

0521 

0521 

0522 

0523 

0524 

0525 

0526 

0527 

0528 

0529 

0531) 

053~ 

0'532 

U533 

0534 

0535 

r 0536 

0537 

0538 

0539 

,- o~o 

0~1 

.' 0542 

0543 

0~44 

0545 

0546 

n547 

0548 

O~' 

0550 

0551 

0552 

0553 

OS:;.' 

0555 

055& 

0557 

~5:5il 

0553 

0560 

0561 

0562 

0563 

0564 

0~65 

0566 

0567 

05£8 

056' 

0571) 

-'( r 

.'-. 

ZDO:ZC1 tZel 

00 110 1:1.,2,-1 

ZL(!) .~1.(J -11 

ZDlm:ZOl(I-lI 

ZOOl!ll 'ZOOI.( H I 

110 aml~U:" 

c 

ZLCIl'Z 

ZOlUl·ZD . 

ZDOl(1) -ZOO 

rm'=CZl)t4.4S*GHOOTI 

ctlTR-S1.S(ll-ZlCll 

EAAOR'PJI;CJl 

EU=O.HF.ROR 

cIlod!fTRt~ 

~dI1fR+O.ltEU 

cH)TR= (SL V( II -ZOLI Lll cH).~TR 

C tH)'cHlTRtO.I21tAA 

C ~'CS3-HtOOI 

COl lID ClWGJ/lG TO Ti[ WITlNOUS PRHHO'JT /,.c..,,,£ 

COl SUBSTI nlTE Elft:J<S fOR THE: on f#) ~G~ 69 

2l-Jan-I9S5 12:09:25 

19-Jan-l~a5 21:06:48 

COl CN L1/£ 99 Tnl, tKJ THE 69 CN LIN!: 296 TO 71 

OU20 If (t .LT. WI GO TO 71 

01125 TYPE 130, t .51,Z,SSI, TG££ , POUT ,HtPO 

Ol1JO rO?J1~T ClJ(,7(FlO.5,X)) 

IU W~#tI 

CtI25 

Ct 

Ct 

Ct 

c* 

Ct 

140 

145 

ISO 

If (t .IT. GRPHTI GO TO 71 

Si( ITJ'SSI 

ACC:~)·Z 

TJH££(ITJ·t 

IT' ITt I 

QRPHT'GRFh-r·GRPH'I 

totHC 

rJ1£fI-I;KMfItItI 

If C(SSNO) .GT. I1IVRO) tll~'RO'(SS2-Z0) 

IF' «(I-TC) .cr. iF) GO TO 150 

GO TO 20 

PRIHTt,'TM 0' ,n!,' £lC .' ,£le,' 2ll .' ,211 

PRIHTt,'T.D. vel·' ,(SSZ-ZD),'TIHE .' ,(t-tll) 

PRIHT*,'HTOUT" ,HTO'JT,' fPLS-' ,FPLS,' RLN NO.:' ,100 

PRIHT.,' 15(21 )=' ,15(21),' FPLSl" ,rPLSI. 

Ct 

Ct 

CCt 

IF' (FltM".EO. G.O) GO TO lEO 

TFltI'TlI+3O 

TFIN=IFIX( tt!) 

WRITE a,15S) TIl, T-TII ,(SSZ-ZO) ,111~,HTO'JT, T()(J1WIX, 

1 Tetm1IH 

155 FDmlT (1X.2CZX,tS.~i,2(tX,r5.1),3(2X,r6.31) 

If (IF1AGS .EO. 0) GO TO 175 

nT-T-T11 . 

VTD=CSS2-Z0) 

A71 

'-""-11 fORT~ V3.5-62 P- 10 

rSDO:[ STtvENS.SHIPSTtlfF]'JPAIc..rOR;17 " 

'. --- ._"-_.-, ~ 

• . 

:1 '. 

1 

.' 

~ 

i 

I 
;., ,- . 

c 21-J~-m5 12:09:35 1.W(-1I fORTRA'l V3 •. H2 Pig. II 

IN~-19~ 21:06:·4G fSOO:(STE\!£NS.SAIPSruffl~'f'AIC .fOR:17 

0571 1'J'tl=IOO 

0572 Tl't1U'ffll o((l'Jft-1)tTmCIW-1)HlT)lJm 

0573 '.ttH I'ffll =C (1Jfl-1) t\ffiC 1'JtI-1)+lJTO I/F.'H 

0574 

0575 IF (I'ffl .£0: 11 GO TO'I65 

~----

057. IF CI'ffl .GT. 21 GO 7() 160 

0577 TSD(l'fflla( (flTl-T1'tt1 ~n IttZ+cflT-ntl(l'J'tl))ttZlttO.5 

0578 VSl)CI'fflI'( (VI01-1.tflCmll Itt2HVTD-'.ttl(r.tIi))tt2Itt11.5 

05n GO TO 165 

Cssg 

0581 160 TSM=C (TSO(IJfl-1)tt2It C I'JtI-ZIHflT-llfl(l'ffllltt2) 

0582 TS:)(r./tII·(TS!W(I'J'tl-1) )ttO.5 

0583 VSMo( (VSI)CIJfl-I)tt2It (11tI-2)HVTD-UtlC 1'J'tl) )tt2) 

0594 VSOCI'ffl)=(VSOAICr.ttl-1I)ttO.5 

0585 

C5S' 165 TlIf\(I'J'tl)=/ffi 

0587 flTloflT ~~ ... ossa 

058' illITE C3,170) rJfl,llflCKttH,TSOCrJfl),U1lCrJfl),VSO(rJfll 

om 170 fO~T (1X,:~,4(2J(,FS.41) 

0591 175 COOItU 

om IF C1ru.G5 .EO. 0) GO TO 180 

0593 6tG=0 

0594 TfIN=fl~TCl) 

0595 IT'l 

OSS CAU. f'LOT3(T1If1,llfl,TSO,U1l,VSD,E£6. TFIH,IT ,~y ,rAY! 

0597 180 r.tOO=l 

" 0598 00 185 1=119:5,795,-10 

0599 lllRTR(r.roJ)=flmT( 11/1000 

0600 KroD=I:DOOH 

0601 185 CC:ITI~ 

060Z !TOT=O.O 

0603 00 IS11 1'1,<0 

0604 ITOT-ITOHISCI) 

0605. 190 CtNTl~,£ 

060& 00 195 1=1,40 

om StOTCI)=FlmTClSC IIl/FlmTCITOT) 

0609 195 coo lt/Jt 

CGO' TfI1P1.2 

0610 TII·O.S 

om 1T=40 

0612 CALL f'LOT2CMTR,StOT ,SlOT, Tll,TflH,1T ,JAY ,rAY, TlI,Zll,ElC, 

0613 7 1:I'A,ISrAS) 

0614 Ct CAU. PlOT1(TlHfE,S1,AC,TII,TFlN,1TI 

om CALL T1MECSlfl 

om CALL MTE(Bt.H) 

0617 

OEI8 illITE 12,200) elf ,BUfF 

om 200 '-cm1T (lX.2CAl211 

Ok"O ' Pl1lJm 

0621 I}() 

A72 

... 

" i 

i' 

.. , .. \ "!P" J _.' ~.i 
o~ 

/ 

21-J;wI-I~ 12:0':35 ~~-\I FCRTPJtI V3.H2 p~ 12 

U-JiII-15SS 21:C6:48 f~:ISTEVM.SH1PSl1ff)VPAI,.F~;17· 

8vtH AtlublllH 

sm PI C COl m lo.. SI\R EXE P.D NtlU lalG 

1£8 PIC COl k'l. lo.. SIf\ tW£ PJ) IMT L(}l(i 

19116 PIC COl REl Lo.. NOSIlR twr P.D WRT LOiS 

25042 

hfl'l'fnctS 

3 

AtlribllltS RtftrrnCH 

~22 m 

Z'3- Z'7 

393- m 

:391- ~l 

394- ~9S 

3SSs m 

'0' 

'. :.,. t 

39" 4% 

146 

56-

234- m m 3:)0 ~1 ~2 ~3 304 

307 308 303 310 ° 311 312 m 331 

332 333 334 335 

295' m 3Q~ 301 302 303 304 309 

310 311 312 332 333 334 335 

397- 4~3 seo 501 502 503 

3922 33l 

3SS' 4~ ~9' 500 501 502 503 

3S9o 4~1 499 m 500 501 502 503 

593- 59~ 

15 61~ m 

16 &1£.\ 618 

Z'7' m 314 

4%- m m 493 505 

m 

3~6' 307 322 325 3<~ 330 :m 340 

491- 4gB 

m o 50Q 505 501 

m- m 314 316 

307· joa 322 325 326 330- 331 337 

340 3~1 (~ 493 500 :01 505 507 

50' 

m- 300(2) 314 316 31B 

A.73 

,/ 

{ 

'i 

-: 

" I 

, _.-. 

'r 

-----r 

/ / , 

;. 

r 

" 

, . \ \ 

, . , 

... It 

t· 

I, 

.' 

2-iJ000484C Rt4 

2-GOOO472S 1t4 

2-iJOOO4674 1*4 

2-GOO04r04 1*4 

2-00004608 1t4 

2-000046OC 1t4 

2-iJOOC4.EO 1*4 

2-GOOoc6C4 1*4 

2-iJOOG4668 1*4· 

2-iJOO046Ca It4 

2-GOGmCC 1t4 

2-00004~: It4 

2-iJOO04670 1*4 

AP-000:0004f I t4 

2-iJoa04714 It4 

2-00C04614 Ri4 

2-iJOC0463a Rt4 

2-GC00463C Rt4 

2-iJOOO4640 RiC 

2-iJOC04G44 R14 

2-o0004S1S Rt, 

2-iJOC0161 C Rt4 

2-0C00462O RH 

2-00004524 Ri4 

2-o00Q4k"3 Rtc 

HO'0462C Ri4 

2-00004ti3fi Rt4 

2-00004634 RU 

2-000048"4 114 

2-00004724 114 

AP-C0000033' It4 

Ai'-OOOOOOOC' 1*4 

AP-Oo~oooce. It4 

2-DOO046t)C 114 

2-00004800 1*4 

2-00004718 1*4 

2-00 CO 4Fii4 114 

AP-00000020' Rt4 

, 

1 

HTP08 3~8-

I 10ba 

114 

516-

6Cb100 33a 

1F1.AG10 no 

IflAGll no 

2S9a 

1fl.AG12 74-

1flAG13 75-

IFl.4G3 68-

(flAGS 27-

IflAG7 ,,-

I FlAGS 70-

IFI.A63 71- mT 29-

I SEAS 3 

IT 100-

ITl 12 

337 

ITlO 12 

1T1l 12 

1Tl2 13 

1T13 13 

IT2 12 

337 

1T3 . 12 

337 

IT4 12 

337 

ITS 12 

337 

IT6 12 

S06 

IT7 12 

ITS 12 

IT3 12 

!TOT toz. J 105-

JAY 3 

JR.'I'f 3 ~ 3 

AAY 28- ~DCO 597-

M..m 101-

rm sn582(2) 

,;> 3 

.:. 

'J :'~~~.~- or ' ~ ~ '~./.~ 

- • _.' • , .a~_ •• 

, 

• ..... :~ ':~a .' .' : ~ ... ... ; ~

2l-JiII-I~SS 12:09:3!I ~1\1(-11 rom,'t/ Vl.5-62 Pi')t 14 

1

IS-Jift-I~..5 21:06:48 rsoo: I STtWlS. SHIPSTurF )VI'A1C.fOR; 11 

3'2 

107 108 10' 110 111 112 113 4 

1160 117 11B 281- 282(2) 283(2) 284(2) 

517(2) 518(21 519(2) ~8o 59' 603- 604 

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ClI1PILATlI!4 STATISTICS 

loft Tilltl 

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*************.*******-***** 

* 

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* 

VIII. APPENDIX B 

* 

* 

* 

*************************** 

B1 

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A. TEST VARIABLES 

The following table gives a listing of the variables used in 

the test matrix. 

Table 81: Fixed-Base Simulation Computer Variables 

and Related Variables and Values 

LIST or YUIAliLlS USEII Ilf nX!lI-lASE SIlI'JIATlO)C 

mI v.\.m.ut~ 

Sea Su •• 0, " 6 

I/U1) I, 1 

TN 1.1, 1.07, 1.05, 1.0), 1.01 

D&ap In I 0.0, -0.2, -0.4 8ec-1 

tD,In. lA, 0.1, 0.7 lee 

lm1 nUll I~ COt!TUTn V.\.qIAlIlt 

Ea!In. lAl 

~ Cole) TAUDIC C .. e) 

--- 0.1 0.3 

0.7 0.7 

Vertical 

Vdoclt, 

Duoplnl! 

Ale C.l.IlI :x Cooe-I ) • C.eC> Cf./I«:2)1°pU opU~'OC) --- --- --- 0.0 - ·O.'S 0.0 

-0.2 S.O ·0.45 0.41 

-0.4 2.5 ·0.4S 0.87 

rnlUST /lin Cgy 

T/II IIAn· Ob) 

-

1.00 

~ 

lo;?4 

~ 

lS7J9 

1.01 ran6 ISS54 

1,03 18422 18193 

1.05 lS011 17847 

1.07 177)] 1751] 

1.10 17%49 17036 

• DUlerent ve1a:htl ar. uled. due to. the c.hana!nl au::!: vtnd •• 

I function of ... Itd.l.lft orchr to .11alnaea tbe effect. of 

the a.rodyn •• ic fore ••• 

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The mean aerodynamic force for a given sea state was determined 

by: 

Where: 

FAZN = WAIT _ FGS 

HKTW 

FAZN - Mean aerodynamic force 

WAIT - Aircraft weight 

HKTW - Scale factor 

FGS - Gross thrust 

(B1) 

The following tablp. shows the values used for the above variables 

as a functiun of sea state. 

Table 82: Variable Values Used in correcting 

for the Mean Aerodynamic Forces 

TRIM FOR T/W • 1.0 

Sea State HKnl WAIT FGS FAlU 

0 0.99046 17036 17530.3 -330.25 

4 0.99046 17036 17353.3 -163.23 

5 0.98734 17249 17509.7 - 39.48 

B. HUD DYNA1UCS 

The following equations describe the hover point dynamics used 

in the IIUD system: 

e = (s.f.) K (x i i ) - K,rel(ThX + x pas pas t on 

+V)+KO ) x 0 long stick 

(Expressions for ey are similar to those for ex.) 

Where: 

s.f. Scale factor 

x = True X position 

B3 

(82) 
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,Kvel ='O~29°/(ft/sec) Ii x = True longitudinal velocity Th .. , . , sec 

(B3) 

n A T,X1 + X n -g T.2 X = s9 

CT,s + 1) X, (T2

S + 1 ) 

(B4) 

Where: 

X .. True longitudinal accelecation 

X = Estimated true acceleration 

x, .. Estimate of high frequency acceleration 

Figure B' shows a graphical definition of the error ex and eye 

Figure B1: HUD Hover Point Dynamics 

c. DATA OUTPUT OF THE PILOTED SIMULATION 

Figure B2 is an example of the output data from the fixed-base 

simulation. Information Used includes the following: 

Page' - All of the title information 

Page 2 - Start and stop times, and T/W information 

Page 3 - Maximum gear z velocity, and the position 

error when the side task (at'titude command 

system) was being flown. 

B4 
· : 

INOTE: 

Since the simulation. was primarily concerned with the vertical 

axis, the highest main gear, designated N (nose) or T (tail), Z 

velocity was used, and the outrigger readings were- ignored. 

L D. PILOTS 

Pilot A - Vernon K. Merrick 

f ' 

Pilot 8 - Glenn G. Ferris 

I. 

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-----------------_._---_ ..... ----.. -~ ... -.---........ _ .. __ ..• 

CVSONG 15:47 JAN 23,'85 RUN <'E'7 PILOT: V. !'ERR ICK 

liAR AVI'A DrHA PR !UTOUT 

F IHAL PR INTJJUT' HOVER APPROACH - VFR TEST 

SHIP tKlTIOt! oil , IISTATE· I,LOH~IIIC.· CLeAt! , ~I[;IGHT. 17S13.0LBS 

L~;;E TURBULEtlCE ON , BACkGROWlv TUj:'B' .(,0 FT.-SEC 

lCSYS • 6 lSU8SYS. 2 II'CASE' 5 IAWB' ~ 

Figure 62: Example Data output from the Piloted Simulation 

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TIre' START AT 

DESCR IPTIott 

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VERT VEL 

)( PILOT ACCEL. 

Y PILOT ACCEL. 

Z PILOT ACCEL. 

PC DEFLECTIOrt 

RC DEFLECT! 011 

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RELATIVE DESCEHT RATE 

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~mER USED· .0 LBS. FUEL USED· 113.J LD5. L~IGHT· 17399.IL8S 

Figure B2, continued: Example Data Output from the Piloteq Simulation 

~.---.~ ~- ... ~ .... ~--.. -,--..... .-._,-'. '-.~ ....... -. -... -........ -;-~-.. _ .. -.. " 

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THE f11HlfUI' f1ISS DISTAHCE L~S LOt~( EC.UAL TO 11.4 FT, 

.--..,. ..--

SHIP ATTITUt:'E 

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Figure 62, continued: Example Da~ Output from the Piloted Simulation 

~HIP VEloel.{ 

:-. • 210~E O;! F-S .•.• • II~iE 02 

:. -. >-H-:C ')1 

1\.:: CCi.f' ;: ''tl 

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E. LISTING or PILOTED SIMULATION RUNS AND RESULTS 

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PILOTED SI~TIOH TEST 'XATRIX 

~'m. TN • 1,10 

~ ... Ill' 

- 1.07 

0.3 0.7 0.3 0.7 

05-8 O· 5-8 

A5-! A 5-! AS-II AS-! 

0.0 OS-A, 10-11 a 5-.A, 10-B 0.0 a 5-A, 10-8 OS-A, S-B 

05-B o 5-B OS-A, 5-B o 5-A, 5-B 

010-B o 5-B 

05-B o 5-B. 

6S-B A 5-B 65-B 65-B 

-o.:! 05-A, 10-B OS-A, 10-B -0.2 o 5-A, 10-B OS-A, 5-5 

Os-a o 6-9 OS-A, S-! o 5-A, S-B 

010-B o S-B 

05-B o 5-B 

AS-B A S-B 65-B 6S-B 

-0.4 OS-A, 10-B OS-A, 10-B -0.4 a 5-A, 5-8 OS-A, S-B 

05-8 o S-B o 5-A, 5-8 OS-A, 5-8 

010-11 o 5-! 

~mg TlW 

- 1.05 

~eng Tl".l 

- 1.03 

0.3 0.7 0.3 0.7 

06-3 0.5-3 

65-! A 5-8 65-8 6S-B 

0.0 OS-A, S-B- OS-A, 5-B 0.0 o IS-A, 10-8 o 11-A, 10-8 

o S-B o 5-B 06-8 05-8 

05-8 05-B 

- 05-B 05-9 

65-B 0. S-B 6 5-B 6S-B 

-0.2 OS-A, 5-B OS-A, S-B -0.2 DIS-A, 15-B DIS-A, lS-B 

OS-If o 5-B o 5-A, 10-B o 5-A, 10-B 

05-8 04-B 

05-8 o 5,.B 

65-1l 0. 5-B 6 S-B 65-B 

-0.4 06-A, 5-B o 5-A, S-B -0.4 o 15-A, 10-B o 10-A, 10-B 

OS-B o 5-B o 5-B o 5-B 

05-B o S-B 

'~eng T/I,I • 1.01 

0.3 0.7 

05-8 o 5-B 

65-8 6 5-B 

0.0 OS-A, 5-8 o 5-A, S-B 

05-A, 5-B o 5-A, 5-B o n-1 S.S.O Hun) V.C. 

05-B 05-8 6 n-1 S.S.4 nun) v.c. 

05-3 o 5-B o n-1 S.S.6 IIUD) v.c. 

65-B A 5-8 o n-t S.S.6 HUDl V.C. 

-0.2 o 5-A, 5-B o 6-A, 6-B o n-1 S.S.6 Hun) ;"C. 

o 5-A, S-B o 5-A, 5.,.B 

05-8 o 5-B n- Number of runs 

1- Pilot 

05-1I o 5-B 

65-8 A 5-B 

-0.4 OS-A, 5-! o 5-A, 5-8 

04-A, S-B OS-A, 5-B 

05-8 05-B 

Figure B3: piloted Simulation Test Matrix 

810 

.

.


,,-.... 

~ n r 

.!- '. PILOTED SIMULATION RUN LISTING Cont • • ! 

U r . . ·SEA STATE 6 Cont • / , 

" • HUD3 : I VELOCITY. COMl-wm SYSTEM ~~1 f. .r4 l.. T/w Z Pilot Run NO.d r.D. Flight P.R. " T '1 max eng w vel ~j Time ~ I 

1,,1 1.05 0.3 0.0 B 621-625 6.2 30.0 6 ~ ! " " 0.2 " 626-630 7.8 32.3 6 • 'I " " 0.4 " 631-6j5 7.0 34.3 5 1/2 it I' " 0.7 0.0 " 636-640 6.4 22.7 6 " 

~~ \.' " " 0.2. " 641-645 9.2 39.8 7 

" " 0.4 " 646-650 7.2 24.1 7 'I l. 1.07 0.3 0.0 A 501-505 8.2 27.9 6 

" " " B 661-665 10.0 16.8 7 

" " 0.2 A 496-500 8.2 21.2 5 1/4 L " " B 656-660 4.5 28.7 3 1/2 

" 0.4 A 491-495 5.5 32.5 4 3/

1, 

" " B 65:1.-655 5.3 22.5 3 1/2 

L. 0.7 0.0 A 476-480 6.2 20.5 51/'}. ... " " B 666-670 8.1 13.1 7 

" 0.2 A 481-485 7.2 17.0 5 1/2 

L " " B 671-675 8.1 20.4 6 .. " 0.4 A 486-490 5.3 23.0 5 (" " " " B 676-680 7.9 34.5 6 

I; 1.10 0.3 0.0 B 696-700 7.6 19.1 7 

" " 0.2 " 701-705 7.0 35.4 5 1/2 

" " 0.4 " 706-710 5.7 38.1 4 

r: " 0.7 0.0 " 69l-695 10.2 22.3 7 1/2 

" " 0.2 " 686-690 7.1 43.4 6 1/2 

" " 0.4 " 681-685 7.3 43.8 5 

L SEA STATE 6 

HUD3 

VELOCITY CO}frlAND SYSTEH 

1.01 0.3 0.0 A 156-160 6.1 33.3 5 

" " " B 186-190 6.5 24.2 6 1/2 

" " 0.2 A 151-155 5.7 34.1 4 1/2 

" " " B 181-185 8.5 34.9 6 1/2 

" " 0.4 A 146-150 5.6 31.2 4 1/4 

" " " B 176-180 9.4 18.3 5 1/2 

" 0.7 0.0 A 171-175 6.7 28.5 5 

" " " B 201-205 6.4 35.4 6 1/2 

" " 0.2 A 166-170 6.5 26.1 5 

" " " B 196-200 7.9 29.6 7 

" " 0,1. A 161-165 7.7 34.2 5 1/2 

" " " 3 191-195 6.7 38.6 5 1/2 

f· 

I. 

r L 8.13 

L 
-~----""'--"'''''·$"--~'''''''",-'~'O~''''~·-·'''·'~-_.'-'''' •• <r"'''~V40.~.l'W'''~~ 

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1 

PILOTED SIMULATION RUN LISTIUG CQnt, 

I: SEA STATE 6 Cont. 

nUD3 

I.- ~ VELOCITY COMMAND SYSTEM 

T/H T Z Pilot Run N·o.s T.D. Flight P.R. 

WCiJC eng w vel Time 

r· 1.03 0.3 0.0 A ~~-1 ~O 7.0 37.7 4 1/2 .1 " " " " 131-135 7.1 28.7 5 

" " " " 271-275 6.7 22.0 5 r· " " " B 291-295 5.5 21.2 5 1/2 

" " " " 556-560 6.0 23.8 5 

" " 0.2 A 91-95 6.6 32.3 4 1/4 

I: " " " " 136-140 4.2 32.2 4 1/2 ; 

I i " " " B 286-290 6.5 26.6 5 

" " " A 401-405 6.5 33.1 5 1/4 

" II· .. B 406-410 5.3· 44.0 5 r: " .. " 551-555 5.1 19.1 6 L: " 0.4 A 86-90 7.4 38.3 4 1/2 

./ " " " 141-145 6.5 24.5 4 1/4 

/ [; " " It 266-270 7.1 42.4 5 1/4 

./ .. " B 281-285 5.4 26.8 4 1/2 

" " " 546-550 7.5 29.8 6 1/2 

(' r' 0.7 0.0 A 111-115 6.5 38.9 4 3/4 

" " It 116-120 6.3 37.1 4 3/4 

.' .. " B 306-310 7.2 17 .1 6 1/2 

" " " 531-535 7.6 23.3 7 

L 0.2 A 106-110 8.0 36.7 4 3/4 

" " " 121-125 6.3 25.6 4 1/2 

" " 276-280 7.7 22.6 5 1/2 

L " B 301-305 5.2 19.9 5 1/2 

" " 536-540 4.7 24.5 5 1/2 j .. I A 101-105 8.1 35.8 4 1/2 I .. " 126-130 7.1 33.7 4 1/2 ! I: It B 296-300 5.8 24.6 5 1/2 It .. 541-545 4.9 25.7 5 I 

I I, 1.05 0.3 0.0 A 216-220 8.9 24.2 5 1/2 

I " " " B 251-255 7.0 30.7 5· 

" .. 0.2 A 211-215 4.4 26.6 4 1/2 

\ : 

" " " B 256-260 5.5 33.6 5 

'-..... .. .. 0.4 A 206-210 6.6 27.7 5 

" .. " B 261-265 7.1 42.4 4 1/2 

f .. 0.7 0.0 A 221-225 7.1 31.S 5 1/2 

t I: .. " " B 246-250 5.0 44.S 6 1/2 

r I .. .. ~.2 A 226-230 7.0 38.3 5 

! " " " B 241-245 8.0 37.2 6 1/2 

\ .-

" It 0.4 A 231-235 6.S 30.9 5 

I " " " B 236-240 9.8 39.9 6 1/2 

L r-- .. 

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J:Wl,;..1...al:;,1..J.-.-..,;. ... .:;, .. ;.:;;;.',;:'".:.:.-:;:,;.: .. ~.:~::.':,.;..'t.;, • ..:.;....;.~$..;.;. ·-"";rt ,... '] ... ..:.,.. ~ .. - •.• i' ""~"'~:;::':_':iF.:-.:::r,.:-=:~ :.'".- j..,:t~. WCIiiCt~ "'::~. ""': _ _ .-~.~ ... :.~ ~'::-. -" ::'...:: •. ~~~.1. ~';';:'~'.'.:::: .::"~'"\~

-. -~~ ..... 

0 ~,!~~8tl!"t":'" :')" , "!%~~t...:..l..._::'"l.:": 1..'-. ' ..... -:-- .. -- ..... _-_. __ . 

'p " 

.. -' 

PILOTED SI~ruLATION RUN LISTING Cont. 

\, 

1.1 SEA STATE 6 Cont. 

HUD3 

/' VELOCITY COHHAND SYSTEM 

I .• T/W

max 

't Z Pilot Run No.s T.D. Flight P.R. eng w vel Time I' 1.05 0.3 0.0 A 216-220 8.9 24.2 5 1/2 " ,I , " " " :13 251-255 7.0 30.7 5 

" It 

, ' 0.2 A 211-215 4.4 26.5 4 1/2 It It It 256~260 5 r.-

B 5.5 33,6 It It 0.4 A 206-210 6.6 27.7 5 It It " B 261-265 7.1 42.4 4 1/2 

I' It 0.7 0.0 A 22l-225 7.0 31.5 5 1/2 

( . It " " B 246-250 5.0 44.5 6 1/2 

" It 0.2 A 226-230 7.n 38.3 5 

" " It B 241-24S' 8.0 37.2 6 1/2 I" 

\ ' 

" It 0.4 A 231-235 6.5 30.9 5 

" " It B 236-240 9.a 39.9 6 1/2 

r 1.07 0.3 0.0 B 386-390 5.0 24.7 4 1/'1. 

(.1 It It " 436-440 5.8 29.5 4 1/2 / " It / A 446-450 5.9 25.2 5 J /4 

1 ' " 0.2 B 391-395 7.8 22.2 4 

\ It It It 441-445 6.1 26.7 3 1/2 It It A 451-455 5.9 31.3 5 It 0.4 B 396-400 5.5 26.2 3 1/2 

{ It A 456-460 6.0 30,.3 4 1/2 i 0.7 0.0 B 381-385 7.4 23.5 5 1/2 It It A 471-475 5.4 39.9 6 

L " 0.2 B 376-380 7.0 27.7 4 1/2 

" It A 466-470 7.4 35.6 5 1/2 

" 0.4 B 371-375 7.1 24.7 4 1/2 ........... It I' " A 461-465 5.7 35.4 4 3/4 

i I 1.10 0.3 0.0 B 321-325 6.7 33.2 3 1/2 I " ", " A 356-360 5.3 32.5 3 3/4 

! i " " " B 721-725 5.2 26.9 4 ! 1- !I It 0.2 " 316-320 5.3 25.8 3 

" " " A 361-365 5.9 29.3 4 " 

! I " It " n 716-720 4.5 52.1 3 1/2 ! " " 0.4 It 311-315 8.2 40.2 3 It It It A 366.-370 6.2 26.2 4 

"j " " It B 711-715 4.7 53.9 2 1/2 I / " 0.7 0.0 " 336-340 6.9 32.6 5 ,/", " " " A 351-355 6.9 34.4 4 1/2 

- " " " B 726-730 5.6 39.0 4 1/2 

L " It 

, 0.2 " 331-335 4.4 29.2 3 ' ~ " " " A 346-350 6.1 35.0 4 ~ ' . " " " B 731-735 5.4 22.1 4 

I"'" ' t It " 0.4 " 326-330 6.4 40.1 3 , I, ! I. " " " 'A 341-345 5.2 48.9 4 '. " It " B 736-740 5.0 33.7 5 1/2 ~ r 

. t 

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0 

P .1' 

PILOTED SIMULATION RUN LISTING Cont, 

1 J ZEA STATE 6 

HUD3 

J.l 

ATTITUDE COMMAND SYSTEM 

T/W

max 

T Z Pilot Run No.s T.D. Flight P.R. eng w vel Time ri 1.01 0.3 0.0 B 1051-1055 8.·7 36.9 8 .. " " 0.2 " 1046-1050 5.7 30.0 7 

" " 0.4 " 1041-104.5 6.6 29.6. 6 

U " 0.7 0.0 " 1056-1060 6.3 25.3 8 

" " 0.2 " 1061-1065 5.9 35.3 6 1/2 

" " 0.4 " 1066-1(}70 6.9 39:6 6 

L 1.05 0.3 0.0 " 1026-1030 9.5 24.8 8 

" " 0.2 " 1031-1035 7.5 29.6 6 1/2 

" " 0.4 " 1036-1040 4.8 35.9 5 1/2 L " 0.7 0.0 " 1021-1025 8.7 27.9 8 

" " 0.2 " 1016~1020 9.9 22.2 7 

" " 0.4 " 1011-1015 6.3 27.3 6 

,-; 1.10 0~3 0.0 " .•. -. 1081-1085 5.0 33.4 5 

" " 0.2 " 1076-1080 6.8 34.8 5 (- " " 0.0 " 1071-1075 5.4 49.2 4 1/2 f[ " 0.7 0.0 " 996-1000 7.3 23.0 7 1/2 

" " 0.2 " 1001-1005 6.7 31.6 6 1/2 

" " 0.4 " 1006-1010 6.7 29.0 6 / 

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ABSTRACT 

The problem of d~termining the vertical axis control 

requirements. for landing a VTOL aircraft on a moving ship deck in 

various sea states is examined. Both a fixed-base piloted 

simulation and a non~piloted simulation were used to determine the 

landing performance as influenced by thrust-to-weight ratio, 

vertical damping, and engine lags. 

The piloted simulation wan ~n using a fixed~base simulator at 

N.A.S.A. Ames Research Conter. simplified vernions of an existing 

AV-BA Harrier modol and an existing head-up display format were 

used. The ship model used was that of a 00963 class destroyer. 

simplified linear models of the pilot, aircraft, ship motion, 

and nhip air-wake turbulence were developed for the non-piloted 

simulation. A unique aspect of the non-piloted simulation was the 

development of a model of the piloting strategy used for shipboard 

landing. ~bis model was .refined during the piloted simUlation until 

it provided a reasonably good representation of' observed pilot· 

behavior. Further refinement could lead to a model suitable for 

prediction of landing performance of VTOL aircraft on ships and as 

the basis of control logic !orautomatic landing. 

A surprising result of this simulation was that, with a good 

station keeping control system and with s.tatistical ship motion 

displayed on the head-up displa~, pilots could consistently perform 

safe landings in sea state 6, with handling q'.1alities that were 

adequate at. thrust-to-weight ratios greater than 1.03 and even. 

i 
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marginally adequate down to thrust-to-weiqht ratios of 1.01. These 

results should' hold quite generally provided that a thrust-to-weight 

ratio' of1 + fl is interpreted as meanir.g that the pilot always has 

the capability of accelerating the aircraft at Aq upward even in the 

presence of ground effect and hot gas reinqestion. 

11 
/ 

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./"1, 
End of Document