The effect of a predictive wind hear chart on pilot landing performance

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April, 1980

THE EFFECT OF A PREDICTIVE WIND SHEAR CHART ON PILOT LANDING PERFORMANCE

• /

by TECHNISCHE HOGESCHOOL DelFT

Eric Nick Solowka

LUCHTVAART- EN RUI. HEVMRTTECHNIEI< BIB lOTHElE •

Kluyverweg 1 - DELFT

1 S-EP. 1980

UTIAS Technica1 Note No. 220 CN ISSN 0082-5263

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THE EFFECT OF A PREDICTIVE WIND SHEAR CHART ON PILOT LANDING PERFORMANCE

by

Eric Nick Solowka

Submitted November, 1979

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Acknowledgement

I would like to thank m:y supervi sor, Dr. L. D. Reid, tor the opp or tuni ty to do this investigation and for his many suggestions in the development of the simulator and associated experiment. Also, I would like to thank Dr. J. H. de Leeuw for his assistance in the final preparation of this technical note in Dr. Reid's absence. Further, I would like to th ank Mr. W. O. Graf for bis valu&ble time in many fruitful discussions and for his assistance in helping me conquer the HP2l00 computer.

I would like to express thanks to the entire UTIAS support staff for their expertise and efficiency.

Finally, I wish to thank the subjects: P. Allen, M. Daniel, Dr. J. H. de Leeuw, A. B. Markov, To Rochford and Dr. P. Sears who donated their time and effort, for their enthusiasm and patience.

This work was supported by the Ontario Graduate Scholarship program and the National Research Council of Canada.

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Summary

A fixed base aircraft simulator ~imited to providing only the longi-tUdinal equations of motion was used to study the effect of a predictive wind shear chart on landing performance through wind shear .

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Contents Acknow1edgement Sumrna.ry INTRODUCTION SIMULATION 2.1 Simulator Hardware 2.1.1 2.1.2 2.1.3 2.1.4 2.1.5 Pilot Controls Instruments Infinity Optics Sound Generation Sta11 Warning 2.2 Computer Simulation 2.2.1 Aircraft Dynamics

2.2.2 Digita1 Program Organization 2.2.3 Generated Display Image 2.3 Simulator Validation

EXPERIMENT

3.1 Effect of Wind Shear on Open Loop Response 3.1.1 Wind Shear Headwind Effect

3.1.2 Wind Shear Tai1wind Effect 3.2 C10sed Loop Response

3.3 Wind Shear Profile

3.4 Wind Shear Predictive Chart 3.5 Experimenta1 Procedure

3.5.1 Subjects

3.5.2 Simulator Fami1iarization 3.5.3 Experimental Program 3.5.4 Eva1uation of Performance RESULTS AND DISCUSSION

CONCLUSIONS REFERENCES BIBLIOGRAPHY APPENDICES TABLES FIGURES iv ii iii 1 1 2 2

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4

5 5 6 7 7

8

8 8 9 9 10 10 10 10 11 12 13 15

16

17

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1. INTRODUCTION

Approach and landing are by far the most onerous phases of flight in terms of pilot work load (effort and concentration) and aircraft performance. Under nor.mal conditions the safety margin associated with these maneuvers is adequate, although under extreme unforeseen or abnormal cond;l.tions the air-craft may deviate significantly from its desired airspeed, attitude and position without hope for recovery. In part this is due to the disappearing tradeoff between altitude and airspeed, as an aircraft operating close to the ground can no longer replace lost airspeed with altitude. The lack of an ideal power plant also adds to the problem since jet turbine and turbo-prop engines require substantial spool-up times and hence the available thrust will lag the required thrust in an emergency. Further, the pilot in his highly attentive state, anticipating the worst of known events that would lead to a balked landing, might miss clues or indications of a more serious abnormal event thereby delaying his intervention and subsequent accident avoidance. The picture drawn then is that landings are safe, as proven by tens of thousands major airlines' operations every day without incident, yet potentially hazardous.

The particular unknown or abnormal condition investigated in this study was the encounter of wind shear on the final approach to landing. Wind

shear, defined as a substantial change of the atmospheric wind's magnitude or direction in a short altitude range, introduces a potential hazard in-herent to the landing operation by unexpectedly causing radical changes in airspeed and inducing flight path deviations. Studies of aircraft open loop response (controls fixed) have concluded that the presence of wind shear during a low level aircraft operation such as a landing, reduced stability and impeded performance (Refs.

5, 7, 8,

10). This poor perfor-mance (i.e., failure to maintain airspeed and desired glide path) , combined with lags as great as two to ten seconds in the detection of shear by the pilot (Ref. 2) can reduce the safety margin in landing to nil as evidenced by several tragic accidents attributed to encounters wi th shear (Refs. 6, 13,14). Shrager (Ref. 12) has further claimed that wind shear has probably

contributed to many other incidents yet remained unidentified as a cause during accident investigation.

Recent advances in acoustic, microwave, and laser Doppler anemometry have made technologically possible the detection of wind shear at and around

airports (Ref. 1). Incorporating all information regarding wind shear detected in this manner into a concise visual cockpit display would reduce wind shear encounters from being potentially hazardous unanticipated events

during the landing approach to predicted or known events and thereby possibly eliminating the hazard. This study investigated the improvement in landing performance through wind shear by the use of a predictive wind shear chart.

2. SIMULATION .

The simulator for the current study was an adaptation of the existing general purpose, fixed base simulation facility at the Institute. In its role as a light transport aircraft it featured a visual display of a runway with glide slope indicators for landing and approach studies. It was equipped with an airspeed indicator, an altimeter, and a power indicator which assisted

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the pilot in re setting the throttle before each experimental run. To enhance realism and improve subject task orientation same nor.mal cockpit sounds were electronically synthesized and infinity optics were added to the visual cathode ray tube display (Fig. 1).

The entire simulation centred on an HP2l00A digital computer which calculated the aircraft position and attitude, generated the required perspective view of the runway, output instrument readings and wind shear values to a Pace TR-48 analog computer, and recorded flight data on magnetic tape.

The mentioned analog computer was programmed with the linearized longi-tudinal equations of motion corresponding to a typical light transport air-craft (Fig. 2). Information regarding the airair-craft's angular and linear velocities was thus made avai1able to the digital computer. The resultant

hybrid analog digital system (Figs.

3,

4) yielded display update r~tes of

25 Hz. Details about the workstation hardware and computer simulation are given in the following sections.

2.1 Simulator Hardware 2.1.1 Pilot Controls

An elevator and aileron control yoke, rudder pedals and throttle were

all present in the pilot's workstation. However, only the throttle and elevator controls we re functional since it was just the longitudinal air-craft-pilot response that was being investigated. The dummy controls were

present,to aid in subject task orientation, i.e., controls that are present

but not functional are not as disturbing to the pilot as missing controls. The elevator aileron assembly (Fig. 5) was salvaged from a defunct Link trainer. The leaf spring that provided the elevator stick force was refitted

so that the resu1tant feel wou1d approximate that of a typical light trans-port aircraft. This was about 11 pounds per inch, push or pull. Elevator yoke movements drove a pulley system that was attached to a potentiometer. The potentiometer output was proportional to the yoke movement and hence gave

a value for elevator angle that was fed into the equations of motion on the analog computer.

The throttle was of conventional design and required about one foot pound of torque to l,llove it in either direction. Again the output was obtained from a potentiometer but since developed engine thrust does not respond instantaneously to throttle movement, the potentiometer output was fed into a simple R-C circuit (Fig. 6)to produce a time lag between throttle position and thrust output. The result was that approximately 75% of the

required thrust, as set by the throttle was available 3 seconds ~fter the

:Lni tial throttle movement , the remaining 25% building up after another 3

seconds (Fig. "7). This lag circuit output was then input to the analog computer.

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2.1.2 Instruments

The power and airspeed indicators were essentially voltmeters cosmetically altered to look like aircraft instruments (Fig.

8).

The power indicator, since it was being driven by the time lag throttle circuit, required no alterations other than overload protection. The airspeed indicator was slightly different in that it was magnetically damped. This prevented needle overswing and instru-ment damage when step changes in the airspeed occurred as was the case when the aircraft penetrated a wind shear layer.

The altimeter was a modified autosyn altimeter that required an alternating current signal to drive it. This was inconvenient since the present simulation was strictly D.C. analoge The A.C. drive was removed and replaced with aservo motor and potentiometer feedback circuit (Fig. 9) which produced an altitude reading praportional to the input voltage from the HP2l00 computer. The rate of climb or descent th at the instrument could follow wi thout ndticeable lag was in excess of 3000 fe et per minute which was more than adequate for a länding simulation.

2.1.3 Infinity Qptics

An inexpensive rectangular Fresnel 2ens was placed between the subjeet's eyes and the display screen (Fig. 10). The benefit was two-fold. First, the introduction of the lens had the effect of increasing the effective window size for thesubject, i.e., the arigle of view both horizontally and vertically was increased for the same size CRT display screen. Second, and more important from a task orientation standpoint was the effect of improving the subjeet's acceptance of the image viewed as part of the real world. The Fresnel lens did this by creating a virtual image that was considerably further away than the screen itself. The eye focusing on this distant virtual image informed the brain that it was (or appeared to be) a distant object thus improving the

simulator realisme

Using a ray diagram (Fig. 11) and fundamental principles of opties the mathematical relationship between the lens characteristics and geometrical

setup yielded an expression for the distanee between the eye and the virtual image created by the Fresnel lens:

D

=

d + f(f - a) a

where D

=

distanee from eye to virtual image, d

=

distanee from eye to the Fresnel lens, f

=

Fresnel lens focal length,

a

=

distanee inside focal plane of display screen.

For small enough values of "a" it can be seen that D can be made very large. Too large a value for D however would be at the expense of image

clarity and definition due to the limited aptical qualities of the inexpensive Fresnel lens. The lens used had a focal length of 12.25 inches so d and a were optimized at 12 and 1 inches respectively producing a virtual image

distance D of 12.5 feet while maintaining an acceptable viewed image (Fig. 12).

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2.1.4 Sound Generation

The siIlD..llated aircraft was turboprop powered, thus it was decided to

have turbine, propeller, and random air.noises present in the cockpit area

during the landing simulation • To acco~lish this, three Texas Instruments

complex sound generation I.C.'s were employed, each chip devoted to the syn-thesis of one sound (Fig. 13) • Since no recordings of the actual noises in a

typical light transport cockpit" were available the pseudo noises were

synthe-sized to sound correct based on the opinion of several exper ienced pilots.

The random air noise, i.e., the noise" in the cockpit that is due to the

air flow passing the airframe was white noise with an upper cut-off frequency. The three dB cut-off frequency with the most promising sound corresponding to

the aircraft equilibrium velocity of 78 knots was about 50 Hz. This sound"

remained fixed in pitch and volume throughout the simulation and wasthus independent of any airspeed perturbations.

The turbine and propeller aerodynamic noises were a combination of an oscillating signal logically "anded" with a white noise signaL The oscilla-ting frequency was voltage controlled and hence could be varied using the same voltage signal that drove the power indicator. The result was that the

prop-eller and turbine RPMvaried from 750 to 2500 and 2000 to 10000 respectively when the throttle was moved from flight idle to maximum RPM.

The three signals were summed via an operational amplifier. This output was then fed into an audio amplifier that drove a speaker attached to the

floqr of the simulator. The speaker was mounted in this manner so that a

small vibration could be felt in the simulator (the major component of which was due to the propeller sound), thereby simulating real airframe vibration. The relative strength of the three noises could be varied by attenuators before they were sunnned so that fine tuning could be do ne • It was discovered that turbine whine in the cockpit was unacceptable for a commercial aircraft so the turbine noise was attenuated such that it could only be heard if a conscious effort were made to notice it.

2.1.5 Stall Warning

The stall warning, although primarily a convenience, added to the overall ability of a subj ect to projeét himself into the task. A simple circuit

(Fig. 14) compared the voltage driving the airspeed indicator with a reference voltage that corresponded to the stall warning speed. When the indicator voltage dropped below reference (corresponding to stall speed) ahorn sounded

and a red light came on in a "push to cancel"

sm

tch on the instrument panel.

The cancel feature was incorporated so that" the stall horn would not become "a

nuisance during the landing flare. The stall warning was reset automatically

about 5 seconds af ter the cancel switch was pushed.

A distinct harn sound was generated by another complex noise generating I.C. The output was through the same audio amplifier used for the normal cockpit sounds.

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2.2 COmputer Simulation 2.2.1 Aircraft Dynamics

The linearized longitudinal equations for the simulation were obtained from Ref.

9

(see Appendix A). The linearization was about a reference equi-librium flight condition specified by the Euler angle Se (Fig. 15) that was determined by the aircraft reference equilibrium velocity Ve, the reference equilibrium wind speed We, and the desired glide slope angle GSA. Values used for the stability derivatives and aircraft parameters were for an equi-librium airspeed of

78

knots and an equilibrium wind velocity of

34

knots. The glide slope ang1e being 15 degrees was chosen as a reasonable extrapola-tion ofcurrent aircraft technology. The desired feature of the equaextrapola-tions used was the explicit appearance of the wind perturbation magnitude term without the wind perturbation rate term as the calculation of the time derivative of the wind perturbation would unnecessarily complicate the simulation. In brief, this was accomplished as follows (the full develop-ment is given in Appendix A) •

Starting with the basic aircraft force-moment matrix equation:

where 6x

=

perturbation of aircraft state,

~u

=

control terms, i.e., elevator and throttle inputs, -c

~W

=

perturbations in wind, i.e., wind shear terms,

and stibstituting in the aircraft pseudo state vector ~ defined by

~=tsx.-Ct:M

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-3-one obtains a matrix equation for the pseudo state of the form

"

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(2-1)

(2-2)

(2-3)

in which the time derivative lM does not appear. The scalar components of this equation were then programmed on the analog computer using standard techniq~es

(Fig. 16). To save space on the analog board, the transformationof the pseudo state back to the normal state tsx. wasleft up to the digital computer.

Estimates for the maximum values of the analog computer variables and the calculated pot set va1ues are given in Tables 1 and 2 respective1y. Numerica~

values for the matrices ~, .Q,1" (g,z + ~.9.3) and

f3

are gi ven at the end of Appendix A.

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2.2.2 Digital Program Organization

The main program written for this investigation was in HP FORTRAN and organized as per the flowchart in Fig. 17 (program lists are given in Appendix B). The two primary functions of the digital computer were to display infor-mation in re.al time and record date, with data reduction and plotting being

done af ter a session of experimental runs was completed •. The program was made

to handle all the transformations required in the six degrees of freedom of aircraft flight for display generation. That is, the perspective view could be calculated for lateral aircraft perturbations as well as longitudinal ones.

The lateral inputs were ·all set to zero however, since the aircraft dynamics

equations were limited to longitudinal motion. .

In order to obtain a real time simulation with display generation occurring as fast as 40 images per second, no complex software functions were employed in

the display calculations • Spe cifi cally, DO LOOPS were avoided by wri ting.· out

operations explicitly. 'Also, no repetition occurred in arithmetic operations.

For example, if the operation A+B occurred more than once, it was assigned to another variable, C say, and C would be used thereafter. Data recording on

magnetic tape reduced the update rate from 40to 25 Hz, but this was still

acceptable as the average human eye only detects display flicker at frequencies below 20 Hz.

The .display generation~ or main body of the program, consisted of the

following sequence (Fig. 17[:

1. sampling the aircraft pseudo state from the analog computer

via a high speed

AID

converter,

2. transforming this pseudo state to the actual aircraft state

vector with respect to the aircraft stability axis,

3.

transforming this state vector to the inertial or earth

reference frame,

4 • . integrating to obtain the aircraft attitude and position

with respect to the runway image (the runway image was

defined in the earth reference frame earlier in the program, i.e., the simulation parameters section),

5. calculation of the perspective view that a pilot would see from the cockpit, the details of which are given in Appendix C.

Knowledge of,the aircraft position in step 4 allowed the altitude to be

output to the altimeter via a multiprogra.mmer

DIA

converter. The .altitude was

also used to read the appropriate wind shear value from a lookup table that was generated in the run parameters section of the program. This wind shear

value was output to the equations of motion on the analog computer by the

DIA

converter, and was also used by the digital computer in transforming the aircraft pseudo state into the actual state.

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Af ter the display coordinates had been generated in step

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they were output to a vector generator th at produced the image on the CRT screen. A test was then made. If the aircraft had not landed then the program continued the loop by re sampling from the analog computer. If the aircraft touched down then the altimeter was reset to 1000 feet and the program returned to the run parameter section where a new wind shear profile could be generated.

Various options in the program were included to assist during experimental runs. To check that the display screen gains and brightness were correct a test pattern was included at the beginning of the program. To assist in

familiarizing subjects with the simulator a freeze display option was available. Data recording and scoring af ter each run was also optional as some runs, the familiarization runs, were not recorded nor scored.

2.2.3 Generated Display Image

The generated display (Fig. 18) consisted of a horizon line, a runway outline,

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sets of peripheral poles, and 3 glide slope T-bar poles.

The horizon line for this study provided the pilot with aircraft pitch information. The runway outline acted as the overall target. The peripheral poles starting at the runway threshold and extending back to the runway mid-point provided same depth of field and height cues, especially when close to the ground, i.e., below 100 feet. The glide slope T-bar poles extending forward 1000 fe et from the, threshold defined the required glide path. That is, an

aircraft following a path such that the top horizontal bars of the poles as viewed by the pilot appeared to lie on top of each other as well as on the end of the runway, then the aircraft was functionally on the desired glide path.

2.3 Simulator Validation

The hybrid analog-digital system was chec~ed by comparing the open loop response of the simulator to a specific wind shear profile with the response of the digital simulation used in Ref. 9 to the same shear profile. To accom-plish this the given equilibrium flight condition flown subject to wind shear profile 70 with the controls frozen at a reference position (see Fig. 23e'for profile 70). The result was that plots of stability axis velocity perturba-tions ~u and ~w in the x and z directions, as well as plots of

e

and q all as functions of time were matched with those obtained from the digital simulation (Figs. 21a, b, c, d). The equilibrium values were the ones to be used during the experimental study.

Although some pilot feedback was obtained regarding the feelof the simulator before the experiment was run, the bulk of pilot opinion was collec-ted during the actual experiments. All pilots agreed that the simulator was realistic and responded as expected. The fact that the "aircraft" did not respond to ,aileron or rudder con trol was not mentioned af ter the familiariza-tion flights by any of the subjects except one, who claimed that he "didn' t miss them at all". One minor complaint was that the random aerodynamic noise did not change its characteristics (pitch and volume) with varying airspeed especially during encounters with wind shear when the aircraft's airspeed noise would most probably be audible, even in a light transport aircraft.

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3 • EXPERIMENT

Wind shear has been defined as the change of the atmospheric wind vector

in a relati vely small amount of space.. The time of transit through a shear

region on descent is thus relatively short leaving only symptoms for the pilot to counter, with little information about the cause. R. S. Bray (Ref. 2) has concluded th at conventional aircraft instrument displays lead to time delays in the perception of wind shear encounters and as such imp ede the pilot's ability to combat the consequences. The present work thus includes a study on what effect complete prior knowledge of the wind shear profile would have on the performance of a pilot in reducing the hazard of an encounter with shear. The current study investigated the landing performance of an aircraft flying through longitudinal wind shear with a vertical gradient. The descriptor

"longitudinal" refers to the direction of the wind vector, i.e., it only has a

single component, parallel to the major axis of the runway. "Vertical" refers

to the space derivative of the wind vector, i.e., the only non-zero component is the one that varies with height, therefore the shear can only be experienced with changes in altitude as in an approach and landing. Information regarding the shear was made available to the pilot before a landing via a predicti ve chart (Fig. 22). Wind shear effects and experimental details are given in the following sections.

3.1 Effect of Wind Shear on Open Loop Response

Any aircraft on encountering longitudinal wind shear will immediately expei'ience an abrupt change in airspeed with little effect on ground speed. This is due to the aircraft's inertia. The resultant unbalanced forces tend to return the aircraft back to its previous equilibrium st.ate. That is, the air speed will re-establish itself for the given thrust setting and provided

th at the new equilibrium -wind vector is different from the initial one, the

ground speed will be the only noticeable difference remaining af ter the

encounter. For a landing aircraft this new ground speed will prcduce an error in the glide path. Two types of longitudinal shear were studied. They are headwind shear and tailwind shear.

3.1.1 Wind Shear Headwind Effect

A wind shear headwind is one that initially increases the relative air-flow and hence airspeed. A landing aircraft encountering this shear headwind experiences a temporary increase in lift which takes the aircraft above the glide slope so that the aircraft is too high with excess airspeed. The asso-ciated increase in drag acts simultaneously to reduce the air speed towards its equilibrium value. This results in ground speed reduction t;Lnd hence a steeper glide path that would take the aircraft ultimately well below the desired glide slope.

3.1.2 Wind Shear Tailwind Effect

A wind shear tailwind is one that decreases the relative airflow. As with the headwind case, there is a sudden change inairspeed on encountering

the shear. The airspeed decreases for the tailwfnd however, wi th a temporary

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loss in lift leading the aircraft below the glide slope with an air speed deficit. The associated decrease in drag allows the airspeed to build back up to its equilibrium value. The ground speed increases as the equilibrium airspeed is attained yielding a shallower glide path than required thus leaving

the aircraft above the glide slope well af ter the transi~nt effects have died

out.

3.2 Closed Loop Response

The pilot untrained in wind shear effects who experiences shear for the first time responds to the transient effects in a manner that aggravates the post transient situation. In the case of the shear headwind, the pilot, when he perceives it, judges the aircraft to be going too fa st and above the desired glide slope. Normally; this would indicate a reduction in throttle setting,

however sinee the 'shear headwind also has t'he effect of ultimately leaving the

aircraft below the require'd giide path without thrust modification, the new

reduced thrust state will lead the aircraft even further below the glide slope. In light of the fact that the engines would require a finite spool-up time this situation is extremely dangerous in that the power required to pull up might not bè available resulting in a crash. The tailwind shear case,

although not as dangerous, has the same effect'of fooling the pilot into an

incorrect ihitial response. That is, for a low and slow aircraft the pilot

increases thrust with the result that the aircraft ends'up too high above .the

glide slope which of ten impedes the completion of a normal landing.

A pilot with sufficient information about the wind shear profile prior to

landing and with adequate training about wind shear -effects might be able to

avoid a go-around or disaster using only conventional methods. That is, by

acting well in advance of the wind shear encounter with appropriate elevator

and throttle response it 'may be possible to reduce the transïent effects

without aggravating the long term res'ult and hence maintaining the desired

equilibrium state.

3.3 Wind Shear Profile

, The wind shear profiles used in the experiment are' gi ven in Figs. 23a to

23f. They were labelled "jet" because there are actually two shear interfaces per profile. The first shear changes the wind from its reference equilibrium value, the second shear re-establishes it thus producing a "layer" or "jet"

during which 'the wind vector is different from its reference equilibrium value.

For all experimental runs the jet layer was between 400 and 600 feet above

ground level. Wind' shear jets that occurred at various other altitudes were

employed during training runs but are not shown.

The jet type shears were employed because they allowed the pilot to regain the reference equilibrium flight condition in most cases. The shear types

were all given a codè number which was used by the computer to generate the

wind shear look-up tables used during the simulation and to prevent subjects fram correlating runs with wind shear types in the event that they accidentally

saw the experimenter's' sheets. The maximum shear magnitude was 1.5 knots per

foot and was used during all training and experiment al runs except phase 7

(discussed later)which employed a milder

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knot p~r foot s~ear.

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3.4 Wind Shear Predictive Ch~t

This· was called a "chart" , rather than a ".display", to avoid confusion

.with the use of the word "display" when referring to the "visual display" of the simulator. In actuality the chart would be a cathode ray tube located in the aircraft cockpit.: Information regarding the shear type and magnitude would be generated using a ground based shear detector, then transmitted to the cockpit where the CRT would be continuously updated •. The chart employed represented a view of such a CRT immediately prior to commencing the approach, and assumed that the wind shear profile was time independent.

The ideal glide slope was represented on the chart by a diagonal line. The presence of shear was indicated by acolour code, wind direction and magnitude was output by digital numbers. The càlour had the advantage of warning the pilot at a glance whether shear was present and at whataltitude. A completely green chart showed that no shear was present • . Wind information was still present, however, for pilot reference. Yellow and red regions on the

chart denoted areas of caution and danger, respectively. The state of the emergency and hence colour were functions of shear strength and altitude. That is, the greater the magnitude of the shear, and the lower it occurred,

the more dangerous it was •. Refer ~o Figs. 24a to 24e for the charts that correspond to the wind shear profiles used in the experiment (note, wind shear profile 10 did not have achart).

3.5 Experimental Procedure 3.5.1 Subjects

A total of six subjects were used in the study. ·All subjects had at least 150 hours of flying experience and five had professional backgrounds other than commercial aviation. Their flying qualifications and eXIE rience are given in Table 3. They were all at least slightly familiar with the dangerous reputation of wind shear.

3.5.2 Simulator Familiarization

Stibjects were all informed that this was a study of the effect of wind shear on landing performance. They were then given a sheet (Table 4) that listed all the pertinent information about the aircraft simulated and the airport for landing. A few trial landings were then done in order to explain the details about the visual display, i.e., the runway peripheral poles were present to give some depth of field and altitude cues close to the ground to allow for anormal visual flare, and the glide slope "T" images extending forward from the runway threshold represented same advanc~d laser landing system and hence were not solid so that the aircraft could fly directly through them without damage. It was :f'urther explained that when the tops of the three landing "T'" s were aligned ad viewed by the pilot, then the aircraft was

correctly positioned on the glide slope. This meant that the pilot's eyes were flying straight down the glide slope rather than the aircraft centre of gravity, however, any error in glide slope control introduced by this slight difference in perspective view was insignificant along that portion of the glide path where flight data IW.er..e being recorded.

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The pilots were instructed to fly as they would normally fly, i.e., throttle governing altitude and elevator governing airspeed. They were not told to do this in any specific order. Prior to each day's session they were all given three free runs during which they were allowed to fly however they pleased. In most cases the ,first run was used as a joy ride, the remaining two being used to practise landing properly. This was so that they could refreshtheir memory of the simulator's handling qualities.

3.5.3

Experimental Program

The experiment al program consisted of eight one-hour sessions (limited to 1 hour per subject per day) that constituted seven phases of the experi-ment. The phases were further divided into training phases and experimental phases. During any session there was a five minute rest period given to the

subjects af ter one half the total number of runs for that session were com-pleted. Each approach and landing required approximately 60 seconds to complete with a 60 second interval being allowed between the termination of one landing and the beginning of another. Subjects were reminded to reset the throttle to the reference equilibrium value before each run. Immediately prior to turning on the display the cockpit sounds were turned on which acted as a final warning and prepared the subject for the appearance of the display.

Af ter each'run during a training phase the subject was told his score so that his behaviour was bëing positively reinforced for good performance. Experimental phase runs, however, were executed with no score feedback so that no learning would occur during the runs when data was being collected. To cancel any effects associated with the order of presentation of the wind

shear profiles all the experimental runs were set up using a 6 x 6 Latin square. Thus each subject flew a different sequence of the three shear profiles during any experimental phase. Subsequent sessions of experimental runs employed permutations of the Latin square so that no pilot flew the same sequence twice.

The details and purpose of each phase are as follows:

Sessions 1

&

2 - Phase 1: Training phase. Total of 20 runs were given each session with no wind shear present, i.e., profile number 10 was used. Each subject obtained consistent performance by the end of the 40 runs such that further runs did not improve their performance. In general they obtained

scores close to the minimum ~alues allowed by the resolution of the glide slope "T" display and airspeed indicat9r (discussed later).

Session

3 -

Phase 2: Experimental phase. Total of 12 runs equally divided between the no shear profile 10, low level jet headwind profile 20, and low level jet tailwind profile 40. The subject was not w~ned of the presence of wind shear. This phase yielded the worst case performance since the subjects were not warned and hence had to fend for themselves.

Session 4,- Phase 3: Training pha,se. Total of 20 runs. The first 10 were similar to low level jet headwind profile 20, the difference being that the shear layer occurred at random heights between 800 and 200 feet. The remaining .10 runs were like low level jet tailwind profile except, as in the previous

case, the jet layer occurred at different altitudes. Although the subject was warned as to the type of shear that would be encountered. in each run, no

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information was given regarding altitude or magnitude. No advice as to tech-niques for countering wind shear was given. Subjects received score feedback af ter each run. This phase was intended to allow subjects to develop their own strategy for combating wind shear so that their run-to-run technique used in Phase

4

would be consistent.

Session

5 -

Phase

4:

Experimental phase. Total of 12 runs given consisting of equal nurnbers of profiles 10, 20 and 40. Subjects were given the wind shear predictive chart before each approach so that they had complete prior knowledge

of the wind shear magnitude, direction and position. No scores were given to the pilots between runs. The outcome of this, phase allowed a conclusion to be drawn about a pilot's ability to cope with wind shear after some experience with it and knowing all the information about it.

Session

6 -

Phase

5:

Training phase. Total of 20 runs. The first ten were of profile 20, the second ten were of profile 40. Subjects were given the predictive chart corresponding to the profile they would he flying before the runs. Prior to the session the effect of wind shear on aircraft flight path and airspeed we re discussed until the subject was completely familiar with the causes and subsequent effects on glide slope and airspeed perturba-tions. A suggestion offered to help cornbát the effect of the shear and

improve glide slope and air speed control was to anticipate the penetration of the shear by leading the aircraft through it. For the headwind encounter this meant reducing throttle before the shear region, then increasing it immediately af ter passing through it, the opposite being valid for the tailwind case., The elevator' was used in the normal manner to control airspeed. Again it was expected that by the end of this phase subjects would have adopted a fixed strategy for combating wind shear and would not change it in the next phase of experimental runs.

Session

7 -

Phase

6:

Experimental phase. Total of 12 runs consisting of

profiles 10, 20 and 40. Subjects had the shear charts and hence total knowledge of shear presence, location and magnitude ~rior to each approach. No score feed-back was given. Comparison of the results of this phase with those of phase

4

allowed the determinat10n of the effect of coaching on subject performance. Session 8 - Phase 7: Experimental phase. Total of 12 runs consisting of equal numbers of ,profiles 60, 70 and 80. In this phase the subjects were divided randomly into two groups of 3 each. One group had the shear prediction charts and the other did not. An assistant was employed for this phase to give all subjects equal attention, but only give the predictive charts to

3

subjects. The experimenter had no knowledge about which subjects were controls and which were given the charts. This was done to prevent the reading in of conclusions in the analysis of results.

3.5.4

Evaluation of Performance

Data points were recorded at the same frequency as the display image update rate, i.e., about 25 HZ. The recorded variables,were the aircraft altitude above ground level, distance from runway threshold, air speed, thrust value, and elevator angle. Recording was ini,tialized at the beginning of the approach and terminated when the aircraft flew below 100 feet AGL. This ,allowed the pilots to flare and land using their own techniques without

altering the results, as down to this point they should have been ideally located on the glide slope.

12

(18)

The altitude vs distance from threshold variables were plotted to show the actual flight path. The remaining variables airspeed, thrust, and elevator angle were plotted vs the altitude so that they could be related to the wind shear profiles flown in order to obtain qualitative information about the subject I s response to the shear. Plots of the experimental runs for Subject 1 are given in Appendix D.

The parameters used in assessing subject performance were the root mean square errors in the airspeed and glide slope angle. A lower bound to these parameters was set by the resolution error of the airspeed indicator and the visual display screen. The lower bound for the RMS airspeed error was 1 knot, and the lower bound RMS glidè slöpe angle was

.3

degrees. Derivation of these bounds is given in Appendix E.

Intra subject RMS errors resulting from the jet headwind andjet tailwind cases of experiment al phases 2,

4

and

6

were compared using the student I S IItll test to a 95% significance level. That is, the conclusions drawn from the IIt 11

test would have a less than 5% probability of occurring purely by chance. Comparing phase 2 with phase

4

allowed a conclusion to be drawn about the

effect practice and the predictive chart had on landing performance. Comparing phase

4

with phase

6

allowed the effect of coaching to be determined. Finally,

comparing phase 3 with phase 6 yielded a conclusion about the effect of practice, coaching and the predictive display on landing performance. The IItll

values are given in Table

5.

Phase 7 allowed an inter-subject investigation to be made as to the effec-tiveness of the entire training program. Herein, subjects were exposed to slightly different wind shear profiles that although not as severe as those used during training still had the same gross characteristics, i.e., a jet headwind and a jet tailwind. More than two samples were being compared so significance of the results was checked using the F statistic to a 95% sig-nificance level as with the IItll

test. Table

6

lists the results.

4 .

RESULTS AND DISCUSSION

.Plots of the runs corresponding to the phase where the subjects encountered shear unexpectedly (Phase 2, see sample in Appendix D) showed the classic res-ponses. Control actuation occurred approximately 25 to 50 feet af ter the shear

encounters. These distances were translated into time delays of 2 to

3

seconds by estimating the average vertical speed through the shear interface to be of the order of 15 feet/sec. The relatively long tîme of 3 seconds was probably due to the lack of instantaneous acceleration clues since the simulator was of thefixed base type and the simulated random airnoise did not vary with airspeed. The 2 second value was in agreement with findings in the literature (Ref. 2). The initial response of all subjects was opposite to the desired long term result, i.e., encountering shear headwind unexpectedly the subjects reduced thrust when it was detected, encountering shear tailwind unexpectedly .. the subjects increased thrust.

The mean scores of each plot of

4

tuns (see samples in Appendix D) were sumnarized by bar graphs in Fig. 26. Any row represents the performance of one subject. The columns are separated first into experimental phases 2, 4, 6

and

7,

and further subdividëd into the three profiles flown during each phase. Inspection of the reference run scores (Fig. 26), that is the runs corre-sponding to the no shear profiles 10 and 60, indicated that the subjects were

(19)

consistently operating close to, and in many cases below the ~n1mum error

allowed due to resolution errors ergo the subjects were all qualified in handling (flying) the simulator. Further, on inspection of the intra-subject errors due to the presence of wind shear profiles 20 and 40 for phases 2, 4 and 6, :it appeared that there was a general trend toward better performance in phase~4 and 6. Ho~ver the "t" values in Table 5 show that the only significant im-provement was in glide slope angle between phases 2 and 6 since all the "t" values in column 5 alone were greater than the required 2.35 for a 95% signi-ficance level. The other columns, having some "t" values below 2.35 indicate that differences between phases 2 and 4, phases 4 and 6, and the difference in air speed between phases 2 and 6 was not significant as there was at least a 5% probability of them occurring by chance. The lack of a significant improve-ment in airspeed control might have been due to the airspeed performance para-meter rather than the pilot's ability since the airspeed fluctuates wildl~ on encountering shear due to the inertia of the aircraft. This fluctuation being somewhat independent of control input thus contributes significantly to the RMS error in airspeed. A better airspeed performance criterion might have been the RMS of the difference between the closed loop and open loop air speed his-tories.

The test for inter-subject variance of phase 7 (Table 6) revealed that the only significant difference in performance between all the subjects (i.e., F statistic greater than 2.77) was in the RMS airspeed error associated with the jet tailwind profile 80. Further inspection of the airspeed error (Phase 7, shear profile 80 of Fig. 26) yielded two groups whose airspeed errors appeared similar, i.e., subjects 1, 3 and 5 making one group, and subjects 2, 4 and 6 making the other. This was confirmed by the F statistics whose value had to be less than 4.26 in order for the scores to be considered similar to within a 5% confidence level. The group of subjects 2, 4 and 6 had the

signi-ficantly better airspeed control to wind shear profile 80. Since it was not known which subjects had the predictive chart, it was initially postulated, on the basis of performance, that the group .of subjects 2, 4 and 6 had the wind shear chart present during the last phase. It was noted however that their performance was inconsistent with the pattern set up by the previous phases. In general, phases 4 and 6 showed that with the predictive chart present the RMS error in airspeed due to the tailwind shear was greater than the RMS error

due to a headwind shear of the same magnitude. Since this trend was due to the learned response af ter many trials it was assumed that the control strategy would remain the same if the chart were present. Thus the initial postulate was revoked and it was concluded that subjects 1, 3 and 5 who exhibited poorer performance in airspeed control to the jet tailwind profile 80 but maintained

a consistent performance trend were the ones who had the predictive shear chart present during phase 7. In fact, this was the case.

, The subjects that obtained the better air speed performance to profile 80 were flying naturally without the aid of the chart. Reviewing the flight history plots of the subjects with the predictive charts showed that they were over-controlling. This might have been due to their anticipation of the

training shears 15 knot magnitude when they only received the milder shear of profile 80. They were, however, told to make note of the shear magnitude whenever they were given a predictive chart where shear was present. In any case this overreaction by the subjects with the display would indicate that before the general use of wind shear predictive displays ean be implemented extensive pilot training programs will have to be designed to fUlly familiarize pilots wi th all aspects of shear, and to develop a sueeessful control strat'egy in combating shear effects.

14

(20)

5 • CONCLUSIONS

The introduction of the predictive wind shear chart alone did not signi-ficantly alter the pilot t s landing performance. Training as to wind shear effect

and limited coaching combined with the presence of the predictive shear display did however improve pilot ability to maintain the aircraft on the desired flight path with little ·effect on airspeed control. This inability to improve airspeed control might not be an indication of poor pilot performance but suggests the adoption of another airspeed performance criterion such as the difference between the open and closed loop response rather than the difference between the closed loop response and the desired response.

The overreaction of pilots to a novel wind shear (that was slightly different

~ from those used in training flights) when the wind shear chart was available to them indicates that although better glide slope control can be achieved, care will have to be exercised wh en introducing a predictive shear display to the general

aviation field. That is, extensi ve training programs regarding pilot technique will have to be developed so that pilots ma.y become familiar with many varieties

of wind shear. These programs should include a definite pilot control strategy, to be developed by future research, that minimizes or removes the hazard of wind shear when encountered on the approach and landing phases of flight.

(21)

1. BeaUlieu, G. 2. Bray, R. S. 3. Etkin, B. 4. Fraser, A. J. 5. Gera, J. 6 . Laynor, W. G. 7. Luers, J. K. Reeves, J. B. 8. McCarthy, J. Bliek, E. Benseh, R. R. Sarabudia, N. R. 9. Reid, L. D. Markov, A. B. Graf, W. O. 10. Sherman, W. L. 11. Shindman, H. D. 12. Shrager, J. 13. 14. REFERENCES

The Effects of Wind Shear on Aircraft F1ight Path and Methods for Remote Sensing and Reporting of Wind

Shear at Airport s • UI'IAS TN" 216, Feb. 1978.

Factors Inf1uencing Toleranee to Wind Shears in Landing Approach. NASA SP-416.

Dynamics of Atmospheric F1ight. John Wi1ey

&

Sons,

Inc.

Deve10pment and Testing of a Fixed-Base Hovercraft Simulator. UI'IAS TN 197, Dec. 1975.

The Inf1uence of Vertica1 Wind Gradients on the Longitudina1 Motion of Airplanes. NASA TN D-6430, Sept. 1971.

A Wind Shear Accident as Evidenced by Information from the Digita1 F1ight Data Recorder. Paper presented at SASI International Seminar, ottawa, Canada, Oct. 1975.

Effect of Shear on Aircraft Landing. NASA CR-2287, July 1973.

Effect of Wind Turbulence and Shear on Landing

Performance of Jet Transports. AIAA Reprint 78-249, Jan. 1978.

The Application of Techniques for Predicting STOL Aircraft Response to Wind Shear and Turbulence

During the Landing Approach. UI'IAS Report No. 215,

June

1977.

A Theoretica1 Ana1ysis of Airp1ane Longitudinal Stabi1ity and Control as Affected by Wind Shear.

NASA TN D~8496, July 1977.

The Eva1uation of a Fixed-Base Hovercraft Simulator. UI'IAS TN 210, May 1977.

The Analysis of Nationa1 Transportation Safety Board Large Fixed-Wing Aircraft Accident/Incident Reports for the Potentia1 Presence of Low-Level Wind Shear. NTIS FAA-RD-77-169, Dec. 1977.

Nationa1 Tran,sportation Safety Board Aircraft Accident

Report: Continental Air1ines Boeing 727-224, N88777, Stap1eton International Airport. Report No. NTSB-ARR-76-l4, May 1976.

Nationa1 Tran~portation Safety Board Aircraft Accident

Report: Eastern Air1ines Boeing 727-225, N88345 E, John F. Kennedy International Airport, Report No.

NTSB-AAR-76-8, March 1976. '

(22)

Gallagher , J. D. Long, M. E.

ROlfe, J. M.

Staples , K. J.

BIBLIOGRAPHY

Wind Shear. Canadian Aviation, Ju1y 1978.

The Air-Safety Cha11enge. National Geographi.c, Vol. 152, No. 2, Aug. 1977.

Keeping Up on the Ground. Aeronautica1 Journal, July 1977.

Current Prob1ems of F1ight Simulators for Research. Aeronautical Journal, Jan. 1978.

Approach and Landing Simulation. AGARD Report No. 632. Simulation • AGARD Conference Proceedings No.

79.

(23)

..

APPENDIX A

LINEARIZED LONGITUDINAL EQUATIONS OF MOTION

The following equations, taken from Ref.

9,

are the linearized equations

of motion:

.

6e = .6q

.

!::,.xl -

-

V _e 6e sin e _e + 6u cos e _e + 6w sin e _e + &1

D.Z

= V _{6e cos e - 6u sin ee}+ l:Jtl _{cos ee} ₊_&3

I e e

& = -!::,.z I

where the aerodynamic force and moment perturbations may be expressed as

!Si = X 6u + X 6w + X 6rr

u w rr

IJl,

=

Z 6u + Z 6w + Z .6w + Z t::,.q + Z 6'1'l + Z 6rr

u w w q 'I'l rr

via the Bryan expansion. The two quant i ties 6'1'l and 6rr are the control variables representing the perturbation in elevator angle (positive for down elevator) and the perturbation in thrust setting respectively.

The dimensional staqility derivativesof the Bryan expansion expressed in terms of known or calculated nondimensional stability derivatives are as follows:

L

(24)

where _ 1 Zw -

2'

t* pV e S C

zei

Z

=

t* q S C q e zq Z

=

q S C T} e ZT} 1 S -c C Mu

=

2'

_PVe mu 1 Mw

= '2

pV_e S

ë

C_na

Mw

=

~

t* pV_e S

ë

Cnä M = t* _qeS

ë

C q mq M = _qeS C C Tt IllTl

-M C

IDcT

GT = _qe S c 7r 7r

a

=

angle of attack of the aircraft zero lift line

.!:!:J.

=

_b.w/Ve

5a

=

angle between thethrust vector and the x axis of F_B

re

=

-Se 1 2-qe

=

'2

_PVe

c

=

mean wing chord

(25)

t* = ë/(2V )

e

L*

=

lift plus the component of thrust in the

lift direction for the reference equilibrium case.

The vector matrix equation corresponding to these equations is:

Alternatively this may be written:

where T t.u = ( t.T) t.7T) -c m 0 0 0 0 0 0 m-Z. 0 w 0 0 0

~

= 0 -M. w Iyy 0 0 0 0 0 0 1 0 0 0 0 0 0 1 0 0 0 0 0 0 1 X _u X _w 0 -mgcos8 _e 0 0 Z _u Z _w mV +zq _e -mgsin8_e 0 0 Mu M M 0 0 0

~

= w q 0 0 1 0 0 0

cos8_e sin8 _e 0 -V sin8 _e _e 0 0

sin8

e -cos8 e 0 V cos8 e e 0 0

(26)

0 X 7r Z Z Tl 7r ~l

=

M Tl M 7r 0 0 0 0 0 0 0 0 0 0 0 0 ~2

=

0 0 1 0 0 -1 -mcos9 msin9 e e -msing_e -mcosg e 0 0 ~3

=

0 0 0 0 0 0 "

-Tc remove the explicit appearance of the time derivative DW define the vector

D:z. :

D:z.

=

!:sx. - C DW

- - 3

-.

Substituting in for ~ one cbtains

(27)

and further

Simplifying yields the final equations:

The actual matrices used in the simulation are given on page A-6. [Note: Matrix BBD

=

(Q2 + ~ Q3)]'

(28)

0001

0002

0003

0004

0005

0006

0007

0008

0009

0010

0011

0012

0013

0014

0015

0016

0017

0018

0019

.

0020

0021

0022

0023

0024

0025

0026

0027

0028

0029

0030

0031

0032

0033

0034

0035

0036

0037

0038

0039

0040

0041

0042

0043

0044

0045

0046

0047

0048

0049

0050

0051

0052

0053

0054

0055

0056

0057

0058

0059

C C C C C C C C C C C C C C C C C APP!H>IX B ~ LIS!'

PROGRAM ACLGF'

VECTOR GENERATOR DISPLAY

SW 8:

CALCULATE SCORE

SW 9:

LAST RUN, REWIND TAPE

SW 10:

KICK OFF TEST PATTERN

SW 11:

RECORD RUN

SW 12;

FREEZE DISPLAY

SW 14;

RESTART BOMBED KEY BOARD

SW 15:

NEW SUBJECT, REWIND TO BEGINNING

IHEAD(1):

FILE NUMBER

SUBJECT NUMBER

IHEAD(2):

IHEAD(3):

IHEAD(4):

IHEAD(5):

PHASE OF EXPERIMENT NUMBER

RUN NUMBER

WIND SHEAR PROFILE CODE NUMBER

DIMENSION W1(135),DW1(134),DATA(10,6)

DIMENSION DTHET(10),DETA(10),DPI(10)

DIMENSION DALT(10),DDOUT(10),DASPD(10)

DIMENSION IHEAD(5),IALPH(5)

COMMON IBUF(8),IBUFA(400),IBUFB(400)

OF RECORD

EGUIVALENCE (DTHET(1),DATA(1,1»,(DETA(1),DATA(1,2»

EGUIVALENCE (DPI(1),DATA(1,3»,(DALT(1),DATA(1,4»

EGUIVALENCE (DDOUT(1),DATA(1,5»,(DASPD(1),DATA(1,6»

EGUIVALENCE (NFILE,IHEAD(1»

C

RESTART BOMBED KEYBOARD-ESPECIALLY WSHPF SUBROUTINE

C

IF(ISSW(14»89,9

9 CONTINUE

CTEST PATTERN

WRITE(2,10)

10 FORMAT(·TEST PATTERN UP·)

CALL BUFA2(IBFA1,IBFA2)

CALL SVECT(IBFA2)

20 IF (ISSW(10»30,21

21 IPT=O

X1=32. X2=224.

CALL VECTC(Xl,Xl,X1,X2,IPT,IBFA1)

CALL VECTC(Xl,X2,X2,X2,IPT,IBFA1)

CALL VECTC(X2,X2,X2,Xl,IPT,IBFA1)

CALL VECTC(X2,Xl,Xl,Xl,IPT,IBFA1)

CALL VECTC(Xl,Xl,X2,X2,IPT,IBFA1)

CALL VECTC(Xl,X2,X2,Xl,IPT,IBFA1)

Xl=O.

X2=127. X3=255.

CALL VECTC(X1,X2,X2,Xl,IPT,IBFA1)

CALL VECTC(X2,X1,X3,X2,IPT,IBFA1)

CALL VECTC(X3,X2,X2,X3,IPT,IBFA1)

CALL VECTC(X2,X3,Xl,X2,IPT,IBFA1)

IBFD=IBFA1

IBFA1=IBFA2

IBFA2=IBFD

DO 25 1=1,20

25 SS=SIN(2./3.)

CALL VECTR(IBFA2,IPT)

r

(29)

C GO TO 20 30 CONTINUE B-2 0060 0061 0062 0063 0064 0065 0066 0067 0068 0069 0070 0071 0072 0073 0074 0075 0076 0077 0078 0079 0080 0081 0082 0083 0084 0085 0086 0087 0088 0089 0090 0091 0092 0093 0094 0095 0096 0097 0098 0099 0100 0101 0102 0103 0104

C**MAKE READY THE MT DATA AGUISITION SYSTEM** C

WRITE(2,40)

40 FORMAT(/'MOUNT DATA TAPEI / )

PAUSE 42 NFILE=l

WRITE(2,43)

43 FORMAT('NEW TAPE? _I) READ(1;44)IALPH

44 FORMAT< 5Al ) C

C CHECK TAPE ON LINE C 45 IF(LOCAL(7»46,48 46 WRITE(2,47) 47 FORMAT('TAPE IN LOCAL I /) PAUSE GO TO 45 48 IF(ISOT(7»51,49 49 WRITE(2,50)

50 FORMAT(ITAPE NOT AT LOAD POINT I /) PAUSE

GO TO 45

51 IF(IALPH(1)-54440B)52,60,52 C

C NOT A NEW TAPE-POS'N AFTER LAST FILE C 52 CONTINUE 53 CALL PTAPE(7,1,0) NFILE=NFILE+l 54 IF(IUNIT(7»54,55 55 CALL PTAPE(7,0,1) 56 IF(IUNIT(7»56,57 57 IF(IEOF(7»58,53 58 CALL PTAPE(7,-1,-1) 59 IF(IUNIT(7»59,60 60 WRITE(2,61)NFILE 61 FORMAT('NEXT FILE: 1,15) JEOF=-1 C C**SIMULATION PARAMETERS** C 0105 RL=3000. 0106 RW=150. 0107 PPW=150. 0108 PLHT=100. 0109 OMD=1000. 0110 DVX=-1.15 0111 DVY=1.75 0112 DVZ=9.00 0113 GSA=15. 0114 PSM=.3 0115 THM=.4 0116 PHM=.3 0117 WRITE(2,65)

0118 65 FORMAT('UPDATE RATE (MSECS)? _I)

(30)

0120 0121 0122 0123 0124 0125 0126 0127 0128 0129 0130 0131 0132 0133 0134 0135 0136 0137 0138 0139 0140 0141 0142 0143 0144 0145 0146 0147 0148 0149 0150 0151 0152 0153 0154 0155 ZlM=80. VBD=O.O Z2M=20. W1M=60. FACT1=2000.*1000./W1M FACT2=-2000./1560. GSAR=.017453*GSA PSMSC=PSM/512. THMSC=THM/512. PHMSC=PHM/512. ETAMS=.6/512. PIMSC=1./512. ZlPMS=ZlM/512./1000. VBPMS=VBD/512./1000. Z2PMS=Z2M/512./1000. IUPDT=-UPDAT RSW=RW/2. PPH=-F'LHT GS1=-OMD/3. GS2=2.*GSl GS3=-OMD HGS1=GS1*TANeGSAR) HGS2=GS2*TANeGSAR) HGS3=GS3*TANeGSAR) WTB1=-HGS:J./4. WTB2=-HGS2/4. WTB3=-HGS3/4. Sl=RL/6. S2=RL/3. S3=RL/2. XOHZ=.5E6 YOHZL=-.3E6 YOHZR=.3E6 AX=306.0 AXN=-306.0 AY=382.5 0156 AYN=-382.5 0157 CGHT=6.0 0158 TT=O.O B-3 0159 CALL BUFA2(IBFA1,IBFA2) 0160 WRITE(2,85)

0161 85 FORMAT(-**SIMULATION PARAMETERS SET**-)

0162 GO TO 880

0:L63 C

0164 C RESTART: BACK SPACE TO LAST FILE MARK

0165 C 0166 86 CONTINUE 0167 CALL KTIME 0168 CALL PTAPE(7,-1,0) 0169 87 IF(IUNIT(7»87,88 0170 88 CONTINUE 0171 JEOF=-1 0172 C

0173 C SUBJECT AND PHASE NUMBER

0:L74 880 WRITE(2,881)

0175 881 FORMAT(-SUBJECT NO.? _-)

0176 READ(1,*)IHEAD(2)

0177 WRITE(2,882)

0178 882 FORMAT(-PHASE OF EXPERIMENT (CODE NO.)? _-)

(31)

C WRITE(2,883) 883 FORMAT(-RUN NUMBER1 READ(1,*)IHEAD(4) --) C**RUN PARAMETERS** C 0180 0181 0182 0183 0184 0185 0186 0187 0188 0189 0190 0191 0192 C 0193 C 0194 C 0195 0196 0197 0198 0199 0200 0201 0202 0203 0204 0205 0206 0207 0208 0209 0210 0211 0212 0213 · 0214 0215 0216 0217 0218 0219 0220 0221 0222 022.3 0224 0225 0226 0227 0228 0229 0230 0231 0232 0233 0234 0235 0236 0237 0238 0239 C C 89 CONTINUE 900 901 CALL KTIME WRITE(2,900)NFILE FORMAT(//-FILE NUMBERt -,15/) WRITE(2,901)IHEAD(4) FORMAT(-RUN NUMBERt -,15//) END OF RUNS,SHUT DOWN1

91 92 920 921 922 IF(ISSW(9»91,922 ENDFILE 7 IF(IUNIT(7»92,920 REWIND 7 WRITE(2,921)

FORMAT(-MOUNT 24K SYSTEM TAPE'-/) PAUSE STOP CONTINUE CHI=O.O CHIR=.0175*CHI TX=-3732. TY=O. TZ=-1000. VE=132. WE=-57. THEQR=-.149803 VEPMS=VE/1000. VNPMS=-VEPMS WEPMS=WE/1000. SCH=SIN(CHIR) SCHN=-SCH CCH=COS(CH!R) STQ=SIN (THEl~R) CTQ=COS(THEQR) TTQ=TAN(THEQR) XINT=TX ZINT=TZ TT=O.O CALL WSHPF(W1,DW1,WE,ISHR) IHEAD(5)=ISHR WP1=W1(101) ICHNO=WP1*FACT1 ICHN1=TZ*FACT2 ICHN2=0 ICHN3=0 IPT=4 JEOF=-1 ITAPE=O

K=-l

CALL SVECT(IBFA1) CALL MPDA(ICHNO,ICHN1,ICHN2,ICHN3,IBFA1) CALL VECTR(IBFA1,IPT)

(32)

0240 0241 0242 0243 0244 0245 0246 0247 0248 0249 0250 0251 0252 0253 0254 0255 0256 0257 0258 0259 0260 0261 0262 0263 0264 0265 0266 0267 0268 0269 0270 0271 0272 0273 0274 0275 0276 0277 0278 0279 0280 0281 0282 0283 0284 0285 0286 0287 0288 0289 0290 0291 0292 0293 0294 0295 0296 0297 0298 0299 C ~5 CALL RECRD(IHEAD,5) DO 923 N=I,50 923 WASTE=SIN(2./3.) C THIS WASTES 50 MSECS.

96 WRITE(2,97)

C

97 FORMAT(-RUN READY-I) PAUSE 2

CALL STIME(IUPDT)

*C**AIC*

POSITION AND ATTITUDE SAMPLING**

C 100 CALL SAMP2 CALL TIME(NPASS) C C FREEZE DISPLAY C C IF(ISSW(12»105,107 105 DO 106 N=1,20 SS=SIN(2./3.) 106 CONTINUE GO TO 510 107 IPT=4

C RECORD DATA SECTION

C

IF(K)150,109

109 IF(TZ+CGHT)110,600 110 IF(JEOF)131,130,150 C

C WRITE END FILE

C C 130 CALL ENDF(7) JEOF=l GO TO 150 131 IF(ISSW(11»132,150 132 ITAPE=ITAPE+l DTHET(ITAPE)=THETA DETA(ITAPE)=ETA DPI(ITAPE)=PI DALT(ITAPE)=-TZ DDOUT(ITAPE)=-TX DASPD(ITAPE)=ARSPD C STOP RECORDING AT 100 FT C IF(-TZ-l00.)135,133 133 IF(ITAPE-l0)150,134 134 CALL RECRD(DATA,120) ITAPE=O GO TO 150 135 IF(ITAPE-l0)1350,137 1350 JJ=ITAPE+l DO 136 J=J.J PlO DO 136 1=1,6 136 DATA(J,I)=O. 137 CALL RECRD(DATA,120) ITAPE=O JEOF=O NFILE=NFILE+l

(33)

0300 0301 0302 0303 . 0304 0305 0306 0307 0308 0309 0310 0311 0312