NASA/TM-97-206299

An Impact-Location Estimation Algorithm

for Subsonic Uninhabited Aircraft

Jeffrey E. Bauer

Dryden Flight Research Center

Edwards, California

Edward Teets

AS&M Inc.

Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273

November 1997

https://ntrs.nasa.gov/search.jsp?R=19980000559 2018-05-11T16:38:54+00:00Z
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ABSTRACT

An impact-location estimation algorithm is being used at the NASA Dryden Flight Research Center

to support range safety for uninhabited aerial vehicle flight tests. The algorithm computes an impact location based on the descent rate, mass, and altitude of the vehicle and current wind information. The

predicted impact location is continuously displayed on the range safety officer's moving map display so

that the flightpath of the vehicle can be routed to avoid ground assets if the flight must be terminated. The

algorithm easily adapts to different vehicle termination techniques and has been shown to be accurate to

the extent required to support range safety for subsonic uninhabited aerial vehicles. This paper describes

how the algorithm functions, how the algorithm is used at NASA Dryden, and how various termination

techniques are handled by the algorithm. Other approaches to predicting the impact location and the reasons why they were not selected for real-time implementation are also discussed.

NOMENCLATURE

Cd

d

F

FTS

Fx

Fy

Fz

g

GRIM

H

ILEA

m

M

RSO

S

T

UAV

V

X

Y

P

0

coefficient of drag

drift distance, ft

force, lbf

flight termination system

force in the x-axis direction, lbf

force in the y-axis direction, Ibf

force in the z-axis direction, lbf

gravity constant, 32.2 ft/sec 2

global real-time interactive map

ground altitude, fl

impact-location estimation algorithm

vehicle mass in its terminated state, slug

wind magnitude, ft/sec

range safety officer

drag reference area, ft 2

descent time, sec

uninhabited aerial vehicle

descent velocity (true velocity), ft/sec

distance traveled to the east, ft

distance traveled to the north, ft

standard day density, slug/ft 3

wind direction, deg
INTRODUCTION

The NASA Dryden Flight Research Center has recently been conducting several programs involving

uninhabited vehicles (fig. 1). In general, these vehicles are designed to operate subsonically at altitudes

higher than 50,000 ft. These vehicles tend to have long wing spans resembling gliders. Most of the vehicles fly with less than 20 lbf/ft 2 wing loading. A variety of propulsion systems are used on these vehicles,

including jet engines, turbo-charged internal combustion engines, and solar electric motors.

These vehicles are currently being flown in restricted airspace over Edwards Air Force Base, California, where NASA Dryden resides as a tenant. History has shown the accident rate of uninhabited vehicles

to be greater than that of piloted aircraft. Consequently, care must be exercised in the operation of such

vehicles to minimize risk to other assets, whether on the ground or in the air. All uninhabited aircraft

operated at NASA Dryden that have the potential of violating range boundaries are required to contain a

flight termination system (FTS).I-3 The purpose of the FTS is to ensure that the vehicle does not leave

the designated airspace or range if normal command and control are lost. The location where the terminated vehicle would impact the ground is also important because the airspace in which these vehicles

EC95 43261-1

(a) The Pathfinder vehicle. (b) The Darkstar vehicle.

EC96 43488-1

EC96 43707-4

(c) The Altus I vehicle. (d) The X-36 vehicle.

Figure 1. Representative UAVs at NASA Dryden that use the ILEA.

EC96 43467-7
(e) The Perseus B vehicle.

Figure 1. Concluded.

EC96 43439-5

operate is above high-value ground assets. A means of predicting where a terminated vehicle will impact

is critical for real-time flightpath management.

Algorithms in use at other range facilities 4' 5 are designed for high-kinetic-energy vehicles, but

NASA Dryden needed an algorithm that accounted for low-speed (low-kinetic-energy) vehicles. The

algorithm would be required to operate in real time, be flexible in its ability to account for a variety of termination techniques, and provide a reasonable estimate of impact location. Therefore, NASA Dryden has

developed and incorporated an algorithm that provides a continuous prediction of where the vehicle will

impact in the event the termination system is activated. Although the actual impact location is important,

the direction of drift in the terminated state can be of more importance. The impact-location estimation

algorithm (ILEA) takes into account vehicle information, including current altitude, its mass when terminated, descent rate, and wind data. The present location and predicted impact location of the vehicle

provide the range safety officer (RSO) with enough information to instruct the uninhabited aerial vehicle

(UAV) operator where to fly in order to minimize the chances of damaging ground assets in the event

of a termination.

The ILEA is one of many techniques available to the RSO for assuring range safety and is integrated

into the NASA Dryden global real-time interactive map (GRIM) display. 6 The GRIM is an advanced

moving map display that provides information such as the present location, altitude, speed, heading, and

range boundaries of the vehicle. For UAV operations, additional operational boundaries, referred to as

UAV keep-out zones (fig. 2), can be displayed. These zones are typically drawn around high-value

ground assets. During the flight, the RSO provides guidance to the vehicle operator regarding the vehicle

flightpath in order to prevent the predicted impact point from entering these regions.

This paper describes the algorithm, how various termination techniques are handled by the algorithm,

and how the algorithm is used at NASA Dryden. Other approaches to predicting the impact location, why

these approaches were not selected for real-time implementation, how the algorithm performed during a

mission, and how the algorithm would have performed in an off-design case are also discussed. The

algorithm has presently been adapted by several UAV operators for use in their command and control

facilities and has been included in the real-time displays by other ranges.
Figure2. UAV keep-outzonesin theEdwardsAFB area.

ASSUMPTIONS AND MODEL

The algorithm is an approximation, so the RSO must provide sufficient margin near assets to account

for the uncertainties of the ILEA. Descent characteristics may not be well known, and emergency situations that require flight termination are often full of surprises. The following assumptions are used in

the model.

• The descent rate of the vehicle in the terminated state is known.

• The vehicle remains intact and descends as a single mass. Components of the aircraft that

descend separately are not modeled. An attempt is made to determine what the most massive component would be, and that piece is modeled. The less massive pieces are assumed to represent less

of a range safety hazard. The impact location is displayed as a discrete point rather than a debris

field, although it is understood that the vehicle pieces will not all impact at the same point. Different vehicles and termination techniques will result in different debris patterns. When large pieces

are expected to result from a termination, the vehicle flightpath is carefully planned to include a

reasonable amount of scatter about the predicted impact point for the unmodeled pieces.

• The vehicle speed or kinetic energy along the flightpath dissipates quickly. Any distance

traveled as a result of the flight speed is not significant and can be ignored. For pieces of the

vehicle that break off or for a destruct-type system, the forward motion is also ignored. Because

such pieces are very unlikely to be dynamically stable, it is assumed that the pieces will have an

effective lift-to-drag ratio of zero. Also, the kinetic energy of any such pieces should be low and

thus not significantly contribute to drift distances.

• The vehicle engines or motors stop operating upon termination.

• Standard day altitude densities from the U.S. Standard Atmosphere, 19627 are used for

the calculations. Although atmospheric conditions are typically different from standard day,

4
analysis has shown that use of the standard day data is reasonable. Actual densities from two days

when missions were conducted were used, and the results were compared to standard day data.

The resulting impact locations differed by approximately 1 percent. The standard day assumption

is reasonable because flight testing of these vehicles typically occurs in good weather. If an

operation is expected to occur in all-weather situations, actual day densities would provide

better results.

• The time for the termination system to activate is small and can be ignored. A finite amount

of time is needed for the termination system to activate and for the vehicle to reach terminal velocity, but these events are not modeled by the ILEA.

• Vehicle midflight weight is used to compute vehicle mass. The exact weight of the vehicle at

termination cannot be determined in advance. If accurate information regarding the vehicle fuel

usage is available, weight could be updated throughout the mission; however, weight updating is

not currently done. Thus, midflight weight for the mission is normally used to compute the

vehicle mass.

• The vehicle is not accelerating in the vertical plane (terminal velocity is assumed for

the descent).

• The vehicle in the terminated condition moves with the air mass (winds).

• The downrange motion of a spinning or uncontrolled vehicle will be small. For vehicles that

descend in an uncontrolled fashion, forward motion can also be ignored. The direction of motion

is reasoned to be random in nature, and the net result will be insignificant when compared to the

effects of drift caused by winds. Some vehicles descend in a spin. Although a spin can be a stable

flight mode with a nonzero lift-to-drag ratio, predicting the heading of any downrange travel is

well beyond current capabilities.

• The winds do not change between updates. Wind data can often be hours old, depending on the

frequency of updates. If winds change substantially between updates, the impact location would

not be accurately predicted.

The ground is modeled as 2372 r, which corresponds to the main runway altitude at Edwards Air

Force Base. The ILEA is based on the model of a vehicle descending under parachute. The descent rate is

thus characterized by a parachute drag coefficient, the parachute reference area, and vehicle mass in its

terminated configuration.

Figure 3 shows the physics modeled. Because the vehicle is assumed to be not accelerating in the vertical plane, the sum of the forces acting on it are zero. Drag is thus equal to weight where

1 V 2

drag = ,_xpx xCd xS (1)

weight = m x g (2)

Wind data can be the largest source of error in the algorithm, Wind data are obtained from weather

balloons launched prior to and throughout the mission. These data are used by the algorithm to predict the

downwind drift during the descent and are contained in a separate file referred to as the "winds file."

In general, data from the balloon are available in 1000-ft increments. Another option for obtaining wind

information involves deriving wind estimates from the UAV onboard navigation systems during

the mission.
Z

(_) Y page into

Assumptions

• Terminal velocity

• Vehicle horizontal velocity = 0

• Drag = 1/2 p x V 2 x Cd x S

• Weight = m x g

Wind

_ Weight

T. Fx = wind component

T. Fy = wind component

_" Fz = 0 => drag = weight

Figure 3. Free-body diagram of modeled system.

970855

Figure 4 shows a simplified flowchart of the code. Beginning at ground level and using an altitude

step size of 1000 ft, an average altitude above the ground is determined. The descent rate at this median

altitude is computed by solving equations (1) and (2) for descent velocity, V. From this computed descent

rate, the time to descend through the altitude band is computed. The wind magnitude and direction at the

center of the altitude band is computed from the data provided in the winds file. The impact location is

repositioned downrange based on the time in the altitude band, magnitude of the wind, and direction of

the wind computed for the current altitude band. The sequence is then repeated by incrementing the altitude until the vehicle altitude is reached.

Figure 5 shows how the impact location is displayed in real time. The predicted impact location is

constantly displayed, even after the termination system is activated, if aircraft altitude data are available.

The ILEA exists in real-time and batch versions. Appendix A shows a copy of one of the batch versions of the ILEA. Inputs required are the weight of the vehicle in its terminated state (mass is computed

in the algorithm from the weight), the equivalent parachute drag coefficient and its associated reference

area, wind magnitude and direction as a function of altitude, and the current altitude of the vehicle. In its

current implementation at NASA Dryden, the real-time version obtains altitude from radar-tracking data.

Other implementations could use telemetered data or altitude transponders from the aircraft.

APPLICATION OF ALGORITHM TO VARIOUS

TERMINATION TECHNIQUES

The algorithm is able to adapt to a variety of termination techniques. Figure 6 shows several common

termination techniques employed by UAVs. The technique shown in figure 6(a), descent under a main
Input

1 H = ground altitude I

@_ UpdateIx, display y)

I Interpolate (p) st density H I

I Compute descent

velocity (V) based on density,

weight, and drag

l

I Compute descent time (1") I T = 1000 flN

I

| B

Interpolate wind direction I

(0) and wind magnitude

(M) st H from wind file I

Compute drift distance (d)

from wind magnitude and d = M x T

descent time

I Sum results I y d sin(e) + y

--[ IncrementH ] H=H+IO00

Figure 4. ILEA code flowchart.

Output

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7
Rogers

Dry Lake

Figure 5. Real-time display of impact location.

970872

970857

(a) Parachute.

Figure 6. Termination techniques.
parachute,is theterminationtechniquefor which theILEA wasoriginally developed.For this technique,

theparachutedragcoefficientandreferenceareaareknownandcanbeusedasinputs.

Figure6(b) showstheuseof a"termination"parachuteto applyanasymmetricdragforceto thevehicle.Theparachuteis of sufficiently smallsizeto allow thevehicleto continueto fly in adescendingturn

to theground.This techniqueis differentfrom thepreviousillustration in that theaerodynamicsurfaces

arestill generatinglift. This typeof terminationtechniquegenerallyresultsin muchslowerdescentrates

thananyof theothertechniques.

Theinputsfor thereal-timealgorithmin this casearederivedby usingthebatchalgorithm.Because

thedescent rate is known, an equivalent coefficient of drag, Cd, and a drag reference area, S, are determined by iteration until the correct descent rate is obtained. Downrange flight (or the distance covered as

a result of the aerodynamic surfaces still flying) is not accounted for because the vehicle is descending in

a constant turn. The predicted impact point does not account for the radius of the turn.

Figure 6(c) shows a termination technique that uses the vehicle control surfaces to put the vehicle into

a spin. This system is modeled in the same manner as the flying descent technique described above.

Although the spinning vehicle may have a nonzero lift-to-drag ratio, the turn radius is small. Predicting

the heading of any downrange travel associated with the nonzero lift-to-drag ratio is well beyond current

capabilities. Such travel is assumed to be random in nature and the net result small. In either case, the

effects of the wind are expected to dominate.

Another common system is a combination termination/recovery system (fig. 6(d)). This type of termination technique is a modification of the technique shown in figure 6(a). Upon receiving the termination

command, a drogue parachute is deployed from the vehicle that puts the vehicle into a rapid descent. At a

predetermined altitude, a larger recovery parachute is deployed. The recovery parachute greatly slows the

(b) Asymmetric drag.

Figure 6. Continued.

970858
(c) Control surface.

970859

(d) Two-stage parachute system.

970860

Figure 6. Concluded.

10
descentsothatlessdamagewill occurto thevehicleupongroundimpact.This systemis handledin two

ways.The first way is to usethe real-timealgorithm,ignoring theeffectsof thedrogueparachuteand

modelingonly themainparachute.Thismodelingis doneby inputtingonly the dragcoefficient andreferenceareaof themainparachute,andzeroingthewind magnitudeabovethealtitudeat which the main

parachutedeploys.An analysisusingthebatchalgorithmshowedamajority of thevehicledrift occurred

underthemain parachute(fig. 7). This approachhasbeenselectedbasedon this analysis.Theseresults

aredependentonthe relativesizesof thetwo parachutesandthealtitude at which the largerparachute

is deployed.

Thesecondway is to usethebatchalgorithmto generateaplot thatincludestheeffectsof bothparachutes(fig. 8). Theplot showsthepredicteddownrangedrift distanceasa functionof altitude.TheRSO

knows the currentaltitude of the vehicleand usesthe plot to determinethe expectedimpact location

relativeto the currentlocationof the vehicle.This approachcanbeusedfor a single-parachutesystem

aswell.

Somevehiclescanbeexpectedto breakupasaresultof anFTSactivation.As discussedearlier,only

themostmassivepieceis modeledin thereal-timeILEA; therefore,thevehicleweight in theterminated

stateis lessthanthe total vehicleweight.Thebatchalgorithmis usedto estimatethe drift of the other

pieces.This estimationprovidesdebris field bounds.That information is usedto carefully plan the

flightpathof thevehicle.TheRSOcanusetheinformationduringaflight to provideanadditionalbuffer

to anyboundaries.

Thereal-timeILEA currentlyusestrackingradarandtheGRIM displayin theNASA Drydencontrol

room.Whenusingthesefacilitiesisnot possible,or in theeventof adisplaymalfunction,thebatchalgorithm is usedto generatea plot of drift distanceasa function of altitude(fig. 8). The RSO canaccess

vehicle altitude and position information from the vehiclecommandand control facility. With this

Altitude,

ft

2.0 x 104

1.8--

1.6--

1.4--

1.2--

1.0--

• 8 --

.6 --

.4 --

0 .1

H = 20,000 ft I

_ain parachute

deployment

I I I I L I I f

.2 .3 .4 .5 .6 .7 .8 .9

Drift distance, nmi

970861

Figure 7. Drift distance with the multiparachute termination technique.

11
1.0 m

North

distance,

nmi

.8--

.6

•4 --

o2 --

0--

- 1.0

- 1.0

I L I 1 I I I I I L

- .8 - .6 - .4 - .2 0 .2 .4 .6 .8 1.0

East distance, nmi

970862

Figure 8. Impact location as a function of termination altitude with the multiparachute technique.

information, predicting the impact location and safely conducting the mission is possible without the

real-time display.

The ILEA is also used in real time for vehicles that follow preprogrammed flightpaths or are autonomous. In these cases, however, options for redirecting the flightpath are limited. The batch ILEA is used

during mission planning before flight to estimate drift distances under various wind conditions, and

buffer zones are applied to all boundaries. The vehicle flightpath is then constructed to avoid those

buffer zones.

PROCEDURAL USE OF THE IMPACT-LOCATION

ESTIMATION ALGORITHM

The RSO is responsible for ensuring the vehicle operation is within the airspace provided for the mission and, in the event of termination, avoiding damage to any ground assets by the vehicle. The real-time

ILEA provides continuous information to the RSO on the predicted impact location of the vehicle if the

FTS is activated.

The purpose of the FFS is to keep the vehicle within the designated airspace or range. The FTS is not

intended to serve as a vehicle recovery system, although some designs serve both functions. If a vehicle

breaks up over the range in an area that does not represent a hazard to any ground assets, the RSO does

not need to activate the FTS. Activation of the system under such conditions, however, may be warranted

in order to lessen the severity of the impact and spread of debris. The advantage the ILEA provides is the

ability to make the termination decision with confidence that the vehicle will not drift into a sensitive

area. In fact, termination can be delayed in order to clear a ground asset. When the FTS has been

activated, the ILEA continues to update the predicted impact point (if accurate altitude data are still

12
available).This updateminimizesthe effectsof uncertainties associated with an assumed descent rate

based on the termination technique and is useful in conducting recovery operations (although little can be

done to affect the impact point).

Because the ILEA provides an estimate of the impact location, the RSO uses judgment to provide

additional separation between the estimated impact location and various boundaries. For example, more

margin would be provided around a boundary that represents a ground asset than around a boundary that

indicates airspace or property limits. Information that would influence the RSO's decision regarding how

much margin to provide includes whether the direction of drift is towards or away from a keep-out zone,

the vehicle altitude, the termination technique, the degree of confidence in the assumed descent rate, and

the time since the last wind update.

DISCUSSION

The algorithm has been adopted by several UAV operators who use it in their command and control

stations. This use provides redundancy with the NASA Dryden control room in the event of a display

failure. The simplicity of the algorithm (approximately 240 lines of code) allows its incorporation with

even modest equipment.

The algorithm has evolved while in use at NASA Dryden. In the first release of the code, an estimate

was made of the 1-g glide distance of the vehicle. A similar display was used in the X-15 program and

was referred to as the energy cardioid. 8 The display was used to show which lakebeds were within

gliding distance in the event of an emergency. For the present application, it was reasoned that if the FFS

system should fail, knowing how far the vehicle could travel would be good. Given the altitudes at which

the aircraft are designed to operate and the fact that most of them resemble sail planes with very high liftto-drag ratios, the predicted glide distances are quite large--in some cases consisting of much of the

western United States. The assumption that the vehicle could glide those long distances was judged to be

unreasonable. If a failure occurs that prevents normal command and control and termination, the vehicle

is unlikely to be able to maintain trim over the speed and altitude band required to reach such far destinations. In addition, the areas were too large to provide useful information to the RSO, so this feature

was eliminated.

Modeling the vehicle equations of motion was initially thought to be required to account for the

kinetic energy of the vehicle along the flightpath. Simulations showed that for vehicles traveling at subsonic speeds, kinetic energy is insignificant when compared to the effects of the wind as the vehicle

descends. In addition, the termination systems tend to rapidly eliminate the forward velocity of the vehicle upon activation. The equations of motion can provide a better estimate for termination techniques that

result in rapid descent rates or for situations where terminal velocity is not reached. Because the systems

modeled result in slow descent rates, terminal velocity is reached quickly, and the complication associated with modeling the equations of motion is not necessary.

Using the equations of motions to rigorously model the vehicle descent is straightforward. With the

level of uncertainty associated with the assumptions and inputs, however, additional accuracy obtained

through such an approach is believed to be insignificant. The added computation requirements may strain

the ability of some systems to perform the necessary computations in a real-time environment, particularly systems that are performing other real-time tasks during the mission.

13
RESULTS

Only one activation of a vehicle FTS has occurred since the ILEA has been in use. The vehicle used a

ballistic parachute termination system (fig. 6(a)). While descending, the vehicle began to break up (its

wing separated because of structural failure) at an altitude of approximately 29,400 ft. At an altitude of

approximately 18,400 ft, the FTS was activated. At an altitude of approximately 16,200 ft, the parachute

blossomed and the vehicle descended to ground impact. At an altitude of 3400 ft, altitude information on

the vehicle was lost because of radar look-angle obstructions.

The actual impact location was approximately 0.6 mi from the impact point predicted when the FTS

system was activated. Debris was scattered along a path approximately 2.5 mi long. The debris furthest

from the main wreckage was the wing section and associated components that departed before FTS activation. The debris field for the pieces under parachute was distributed along a line approximately 150 ft

in length where the vehicle was dragged along the ground. The ILEA did its job well in that it informed

the RSO that no ground assets were in the area and none were in the direction of drift.

The discrepancy in impact location can be attributed to a number of factors. The primary contributor

is the result of a change in the winds. Had the best available wind information been included in the algorithm, the predicted impact point would have been less than 0.1 mi from the actual impact point. More

accurate wind information was available only after the accident as a result of the detailed accident

investigation that ensued. At the time of the accident, wind updates were generally not made during

the mission.

Because the vehicle broke up, the weight of the vehicle under parachute was different from that

assumed by the model. Despite the weight differences, the actual descent rate closely matched the

assumed descent rate. The vehicle was in a vertical descent when the FTS was activated. This vertical

descent is consistent with the assumption that the forward velocity of the vehicle will dissipate quickly

and not influence the final results in a significant way.

After the successful utilization of the algorithm in the above situation, the algorithm was tested to see

how it would perform in an off-design case. The case selected was an accident involving the X-31

research aircraft. The X-31 aircraft is a full-scale, piloted, fighter-class vehicle that uses thrust vectoring

for agility at high-angle-of-attack flight conditions. 9 The vehicle weighs approximately 16,500 lbf and

has a cruise speed of 300 kn.

The X-31 case violates several assumptions used in developing the model for the ILEA. The X-31

postdeparture dynamics were not expected to match the UAV model. In spite of that fact, the drift direction was accurately predicted and after correcting for "non-UAV, post-FTS" behavior, the impact point

was predicted within 0.25 mi. Appendix B provides a detailed discussion of this case.

CONCLUDING REMARKS

An impact-location estimation algorithm (ILEA) has been developed and is in use at NASA Dryden

Flight Research Center that estimates the impact location for subsonic (low-energy) uninhabited aerial

vehicles. The algorithm accounts for the most recently observed winds and the descent rate of the vehicle.

The ILEA has the flexibility to account for a variety of termination techniques and is computationally

simple. The algorithm operates well in the real-time flight test environment.

14
Despitethesimplistic modelusedby theILEA, theresultsareadequateto performthe rangesafety

function in anareathatcontainshigh-valuegroundassetsthat mustbe avoidedin theeventof termination. Thealgorithmhasbeenusedsuccessfullyin anactualflight termination,predictingtheimpactpoint

to within 0.6 mi of the actualimpactlocation.TheILEA hasbeenadaptedby severaluninhabitedaerial

vehicleoperatorsfor usein their commandandcontrol facilities andhasbeenincludedin real-timedisplaysby otherranges.

Dryden Flight Research Center

National Aeronautics and Space Administration

Edwards, California, August 25, 1997

15
APPENDIX A

FORTRAN CODE

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The code for one of the batch versions of the impact algorithm is as follows.

PROGRAM Traj9

CALCULATES IMPACT POINT AS A FUNCTION OF ALTITUDE,

AIRSPEED, AND HEADING FOR A GIVEN ALTITUDE

Originally written by Brent Cobleigh.

Modified by Elise Gravance and Jeff Bauer.

Last modified December 3, 1996 by Jeff Bauer.

DIMENSION VX(100),VY(100),

WINDMAG( 100),WINDDIR( 100),HVT( 11 ),

ALT( 100),DIR( 100),DENSITY( 100),CHDTIME(100),

SPD( 100),ALTH( 100),XHDOT( 100),RHORTO(100)

COMMON/DATA/ICOUNT,ALTDOT,DIRECTION,XMAG,RHO,

ALTH,XHDOT,CTRALT,ALT,SPD,RHORTO,DIR,

WEIGHT,CHALTDOT,SCD,VVCURR,DENSITY,PLD,

VVLOW,VVHIGH,BPALT,CURDENS,CURALT,

HVT

AIR DENSITY* 1000 FROM ALTITUDE = 0 TO 99,000 FT IN 1000-FT

INCREMENTS

DATA DENSITY/2.3769,2.3081,2.2409,2.1752,2.1110,2.0482,

1.9869,1.9270,1.8685,1.8113,1.7556,1.7011,1.6480,

1.5961,1.5455,1.4962,1.4480,1.4011,1.3553,1.3107,

1.2673,1.2249,1.1836,1.1435,1.1043,1.0663,1.0292,

.99311,.95801,.92387,.89068,.85841,.82704,.79656,

.76696,.73820,.71028,.67800,.64629,.61608,.58727,

.55982,.53365,.50871,.48493,.46227,.44067,.42008,

.40045,.38175,.36391,.34692,.33072,.31527,.30055,

.28652,.27314,.26039,.24824,.23665,.22561,.21508,

.20505,.19548,.18336,.17767,.16938,.16148,.15395,

.14678,.13993,.13341,.12719,.12126,.11561,.11022,

.10258,.09770,.09306,.08865,.08445,.08046,.07665,

.07304,.06960,.06632,.06321,.06024,.05742,.05473,

.05217,.04974,.04742,.04521,.04311,.04111,.03920,

.03739,.03566,.03401/

DRAG CHUTE PARAMETER SCD=(CHUTE AREA)(CHUTE DRAG COEFFICIENT)

Open output files

OPEN(UNIT=7,FILE='output',status='unknown')

OPEN(UNIT= 11,FILE='crash',STATUS='UNKNOWN')

READ IN CURRENT WINDS FROM FILE WIND.DAT

CALL READWIND

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INPUT CURRENTALTITUDE, CALIBRATED AIRSPEED,HEADING,

GROUNDSPEED,WEIGHT, AND CdS

PRINT*,'INPUTAIRCRAFT ALTITUDE (FT)'

READ(*,*) CURALT

PRINT*,'INPUTAIRCRAFT WEIGHT (LB)'

READ(*,*) WEIGHT

rmass= weight/32.174

PRINT*,'INPUTAIRCRAFT CdS(FT^2)'

READ(*,*) SCD

write(7,*)'Alt, Airspd,Heading,GroundSpd,Weight,mass,SCd'

write(7,*)curalt,vt,theta,vg,weight,rmass,scd

CALCULATE TIME TO PASSTHROUGHEACH ALTITUDE BAND

USEALTITUDE BANDS OF 1000FT

GRDALT=2372. ! GROUNDALTITUDE AT EDWARDS

ALT I=GRDALT ! ALTITUDE BREAKPOINT

ALT2=3000. ! ALTITUDE BREAKPOINT

ICOUNT=0

Time=0.0 !Time to impactfrom altitude

DO WHILE (ALT 1.LT.CURALT)

ICOUNT=ICOUNT+1 ! DUMMY COUNTER

DALT=ALT2-ALT 1

GET HDOT FORTHE CENTEROFTHE CURRENTALTITUDE BAND

CTRALT=(ALT2-ALT 1)/2 + ALT l

write(7,*)'CTRALT =',ctralt

! CENTEROF ALTITUDE BAND

CALL HDOT !INTERPOLATEHDOT

CHDTIME(ICOUNT)=DALT/CHALTDOT ! CALCULATE TIME IN Alt BAND

TIME=Time + chdtime(icount)

CALL WINDD ! INTERPOLATEWIND DIRECTION

WINDDIR(ICOUNT)=DIRECTION

CALL WINDM ! INTERPOLATEWIND MAGNITUDE

WINDMAG(ICOUNT)=XMAG* 1.69 ! CONVERTMAGNITUDE TO FT/SEC

INCREMENT ALTITUDE BAND BY I000 FEET

ALT 1=ALT2

ALT2=ALT2+ 1000.

IF (ALT2.GT.CURALT) THEN

ALT2=CURALT

ENDIF

END DO

COMPUTEIMPACT POINT

DO 200J=1,ICOUNT

CALCULATE X & Y GROUNDTRACK VELOCITY ONLY SUBJECTTO WINDS ALOFT

17
C

C

C

C

XWINDHD=(270.-WINDDIR(J)) !WIND HEADING

VX(J)=WINDMAG(J)*COS(XWINDHD/57.3)

VY(J)=WlNDMAG(J)*SIN(XWINDHD/57.3)

200CONTINUE

CALCULATE X & Y GROUNDTRACK DISTANCEFROM START

XDIST=0.

YDIST=0.

C

C

C

C

C

C

DO 210K=I,ICOUNT

XDIST=XDIST+VX(K)*CHDTIME(K)

YDIST=YDIST+VY(K)*CHDTIME(K)

10CONTINUE

CALCULATE TOTAL DISTANCETRAVELLED

DISTANCE=SQRT(XDIST**2+YDIST**2)

WRITE RESULTSTO OUPUTFILES

write(7,*)'Time =',time

write(11,*)'East(ft), North(ft), TotalDistance(ft)'

WRITE (11,140)XDIST,YDIST,DISTANCE

140FORMAT (3(F12.2,','))

CLOSE(7)

CLOSE(11)

STOP

END

************** _**:_********* _******************************:_***********

SUBROUTINE HDOT

C

C

C

C

C

C

C

CALCULATES CHUTE HDOT AT GIVEN ALTITUDE

DIMENSION HVT(11),

ALT( 100),DIR( IO0),DENSITY(100),

SPD( 100),ALTH ( 100),XHDOT( 100),RHORTO(100)

COMMON/DATA/ICOUNT,ALTDOT,DIRECTION,XMAG,RHO,

ALTH,XHDOT,CTRALT,ALT,SPD,RHORTO,DIR,

WEIGHT,CHALTDOT,SCD,VVCURR,DENSITY,PLD,

VVLOW,VVHIGH,BPALT,CURDENS,CURALT,

HVT

INTERPOLATE DENSITY AT CURRENT ALT

DO 10 I=1,100

DUMALT2=REAL(I) * 1000.

DUMALT 1=REAL(I- 1)* 1000.

IF ((DUMALT2.GT.CTRALT).AND.(CTRALT.GE.DUMALT 1)) THEN

CURDENS ITY=(CTRALT-DUMALT 1)/(DUMALT2-DUMALT 1 )*

(DENSITY(I+ 1)-DENSITY(I))+DENS ITY(I)

CURDENS=CURDENS ITY/1000.

GOTO 30

18
C

ENDIF

10CONTINUE

C

C CALCULATE CHUTEALTITUDE RATEFROM EQUATION QBAR*S*CD=WEIGHT

C WHEREQBAR=.5*CURDENS*V**2

C

30Continue

CHALTDOT=SQRT(2*WEIGHT/CURDENS/SCD)

RETURN

END

SUBROUTINEREADWIND

C

C

C

READ IN CURRENT WINDS FILE

DIMENSION HVT( 11 ),

ALT( 100),DIR( 100),DENS ITY (100),

SPD( 100),ALTH( 100),XHDOT( 100),RHORTO(100)

COMMON/DATA/ICOUNT,ALTDOT,DIRECTION,XMAG,RHO,

ALTH,XHDOT,CTRALT,ALT,SPD,RHORTO,DIR,

WEIGHT,CHALTDOT,SCD,VVCURR,DENSITY,PLD,

VVLOW,VVHIGH,BPALT,CURDENS,CURALT,

HVT

OPEN(UNIT=8,FILE='wind.dat',STATUS='OLD')

DO 20 I=1,100

READ(8,*) ALT(I),DIR(I),SPD(I),dens

IF(ALT(I).LT, 1) THEN

GOTO 30

ENDIF

20 CONTINUE

30 CLOSE(8)

END

SUBROUTINE WINDD

C

C

C

INTERPOLATES WIND DIRECTION FROM FILE

DIMENSION HVT(11),

ALT( 100),DIR( 100),DEN SITY(100),

SPD( 100),ALTH ( 100),XHDOT( 100),RHORTO(100)

COMMON/DATA/ICOUNT,ALTDOT,DIRECTION,XMAG,RHO,

ALTH,XHDOT,CTRALT,ALT,SPD,RHORTO,DIR,

WEIGHT,CHALTDOT,SCD,VVCURR,DENSITY,PLD,

VVLOW,VVHIGH,BPALT,CURDENS,CURALT,

HVT

DO 10 I=1,100

IF ((ALT(I+ 1).GT.CTRALT).AND.(CTRALT.GE.ALT(I))) THEN

IF (ABS(DIR(I+ 1)-DIR(I)).LT.90.) THEN

DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(DIR(I+ 1)-DIR(I))+DIR(I)

ELSEIF ((DIR(I+ 1).LE. 180.).AND.(DIR(I).LE. 180.)) THEN

19
DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(DIN(I+ 1)-DIN(I))+DIN(I)

ELSEIF((DIN(I+ 1).LE.270.).AND.(DIN(I).LE.270.).AND.

(DIN(I+ 1).GT.90.).AND.(DIR(I).GT.90.))THEN

DIRECTION=(CTRALT-ALT(I))/(ALT(I+I)-ALT(I))*

(DIN(I+ 1)-DIR(I))+DIR(I)

i_LSEIF ((DIR(I+l).LE.360.).AND.(DIR(I).LE.360.).AND.

(DIN(I+ 1).GT. 180.).AND.(DIR(I).GT. 180.)) THEN

DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(DIR(I+ I)-DIR(1))+DIR(1)

ELSEIF ((DIN(I+ 1).LE.360.).AND.(DIR(I+ 1).GT.270.).AND.

(DIR(I).GE.0.).AND.(DIR(I).LT.90.)) THEN

• DUMDIR=DIR(I)+360.

DIRECTION=(CTRALT-ALT(I))/(ALT(I+I)-ALT(I))*

(DIN(I+ 1)-DUMDIR)+DUMDIR

• IF(DIRECTION.GE.360.) DIRECTION=DIRECTION-360.

ELSEIF ((DIR(I).LE.360.).AND.(DIR(I).GT.270.).AND.

(DIN(I+ 1).GE.0.).AND. (DIN(I+ 1).LT.90.)) THEN

• DUMDIR=D1R(I+ 1)+360.

DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(DUMDIR-DIN(I))+DIR(I)

• IF(DIRECTION.GE.360.) DIRECTION=DIRECTION-360.

ELSEIF ((DIR(I).LE.270.).AND.(DIR(I).GT. 180.).AND.

(DIR(I+ 1).GE.0.).AND.(DIR(I+ 1).LT.90.)) THEN

IF ((DIR(I)-DIR(I+ 1)).LE. 180.) THEN

DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(DIR(I+ 1)-DIR(I))+DIR(I)

ELSE

DUMDIR-DIR(I+ 1)+360.

DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(DUMDIR-DIR(I))+DIR(I)

ENDIF

ELSEIF ((DIR(I+ 1).LE.270.).AND. (DIR(I+ 1).GT. 180.).AND.

(DIR(I).GE.0.).AND.(DIR(I).LT.90.)) THEN

IF ((DIR(I+ 1)-DIR(I) ).LE. 180. ) THEN

DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(DIR(I+ 1)-DIR(I))+DIR(I)

ELSE

DUMDIR=DIR(I)+360.

DIRECTION=(CTRALT-ALT(I))/(ALT (I+ 1)-ALT(I))*

(DIR(I+ 1)-DUMDIR)+DUMDIR

ENDIF

ELSEIF ((DIR(I).LE.360.).AND.(DIR(I).GT.270.).AND.

(DIR(I+ 1).GE.90.).AND.(DIR(I+ 1).LT. 180.)) THEN

IF ((DIR(I)-DIR(I+ 1)).LE. 180.) THEN

DIRECTION=(CTRALT-ALT(I))/(ALT(I+I)-ALT(I))*

(DIR(I+ 1)-DIR(I))+DIR(I)

ELSE

DUMDIR=DIR(I+ 1)+360•

DIRECTION=(CTRALT-ALT(I))/(ALT(I+I)-ALT(I))*

(DUMDIR-DIR(I))+DIR(I)

IF (DIRECTION.GE.360.) THEN

20
DIRECTION=DIRECTION-360.

ENDW

ENDIF

ELSEIF((DIR(I+ 1).LE.360.).AND.(DIR(I+1).GT.270.).AND.

(DIR(I).GE,90.).AND.(DIR(I).LT.180.))THEN

IF ((DIR(I+ 1)-DIR(I)).LE. 180.)THEN

DIRECTION=(CTRALT-ALT(I))/(ALT(I+I)-ALT(I))*

(DIR(I+ 1)-DIR(I))+DIR(I)

ELSE

DUMDIR=DIR(I)+360.

DIRECTION=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

• (DIR(I+ 1)-DUMDIR)+DUMDIR

IF (DIRECTION.GE.360.) THEN

DIRECTION=DIRECTION-360.

ENDIF

ENDIF

ENDIF

GOTO 20

ENDIF

10 CONTINUE

20 RETURN

END

SUBROUTINE WINDM

C

C

C

C

C

INTERPOLATES WIND MAGNITUDES

DIMENSION HVT(11),

ALT( 100),DIR( 100),DENSITY (100),

SPD(100),ALTH(100),XHDOT(100),RHORTO(100)

COMMON/DATA/ICOUNT,ALTDOT,DIRECTION,XMAG,RHO,

ALTH,XHDOT,CTRALT,ALT,SPD,RHORTO,DIR,

WEIGHT,CHALTDOT,SCD,VVCURR,DENSITY,PLD,

VVLOW,VVHIGH,BPALT,CURDENS,CURALT,

HVT

INTERPOLATE WIND MAGNITUDE

DO 10 I=1,100

IF ((ALT(I+ 1).GT.CTRALT).AND.(CTRALT.GE.ALT(I))) THEN

XMAG=(CTRALT-ALT(I))/(ALT(I+ 1)-ALT(I))*

(SPD(I+ 1)-SPD(I))+SPD(I)

GOTO 21

ENDIF

I 0 CONTINUE

21 RETURN

END

21
APPENDIX B

X-31 AIRCRAFT DISCUSSION

On January 19, 1995, as the X-31 was returning from a research mission and descending at an altitude

of 20,000 ft, a failure of the airdata system caused the vehicle to depart controlled flight and crash. Data

from this accident were replayed through the real-time impact-location estimation algorithm (ILEA) to

determine how well the impact point would have been predicted. The uncorrected results were disappointing, largely because of the nature of the departure.

For purposes of this investigation, the wind information used was obtained from a weather balloon

launched approximately one-half hour after the accident. The winds for this case were very low, and

errors in the prediction were not likely to have been a result of inaccurate wind data. Pilot ejection was

considered as the flight termination system activation point. Because a prediction of the vehicle descent

rate was not available beforehand, a descent rate was obtained by dividing the altitude loss (from ejection

to impact) by the time from ejection to impact.

Ejection occurred at an altitude of approximately 20,500 fl above sea level. From this point, the ILEA

predicted the impact would occur 597 ft to the north and 557 ft to the west of the ejection location. In

actuality, the vehicle impacted approximately 4083 ft to the north and 2947 fl to the west. The primary

contributor to this error is the postdeparture behavior of the aircraft.

During this departure, the vehicle gained approximately 800 ft in altitude before entering into altitude

excursions varying as large as +100 ft. The vehicle airspeed significantly decreased and the pilot ejected.

As a result of the reduced airspeed, the vehicle righted itself following the pilot ejection and "flew" for

approximately 20 sec in a north-northwestern direction. The vehicle continued to descend and gradually

increase airspeed until it once again departed. This second departure resulted in a significantly larger altitude excursion than the previous one. The vehicle never recovered from this departure, and altitude

excursions continued through the final descent. The engine continued to operate until impact.

This sequence violated two basic assumptions of the ILEA, the most significant being the approximately 20 sec of "flight" where the vehicle covered approximately 2430 ft to the north and 1470 ft to the

west. This "flight" accounts for the majority of the error between the predicted and actual impact locations. The second assumption violated was that the vehicle would not accelerate in the vertical plane

during the descent. Because the vehicle actually increased altitude during portions of the descent, this

assumption was violated. When the "flight" distance between departures is subtracted, the drift direction

was accurately predicted and the location was predicted to within approximately 0.25 mi.

22
REFERENCES

1NASA Dryden Flight Research Center, Basic Operations Manual, Rev. 5, Feb. 1995.

2United States Air Force Range Commanders Council, Flight Termination Systems Commonalio,

Standard, Standard 319-92, 1992.

3United States Naval Air Warfare Center, "Requirements for Flight Termination Systems," Weapons

Instruction, NAWCWPNSINST 5100.2, 1993, Appendix B.

4Beyma, Robert J., Flight Safety Range Safety Officer Training Manual, NASA/Wallops Flight Facility,

Sept. 1993.

5Balach, Dean, Flight Safety Programs and Enhancements for Sounding Rockets at Wallops Flight

Facility, NASA TM- 111284, 1995.

6Comperini, Robert and Donald C. Rhea, "Development of an Interactive Real-Time Graphics System

for the Display of Vehicle Space Positioning," AIAA-88-2167, May 1988.

7National Aeronautics and Space Administration, United States Air Force, United States Weather

Bureau, U. S. Standard Atmosphere, 1962: ICAO Standard Atmosphere to 20 Kilometers, Proposed

ICAO Extension to 32 Kilometers, Tables and Data to 700 Kilometers, 1962.

8Cockayne, William G., Description of an Energy Management System for the X-15, NASA

CR-96006, 1968.

'_Herbst, W. B., "X-31A At First Flight," Flying Qualities, AGARD-CP-508, Feb. 1991, pp. 29-1-29-8.

23
REPORT DOCUMENTATION PAGE FormApproved OMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing irlstructlons, search=rig existing data sources, gathering and

maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of informat_on.

including suggestions/or reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204. Arlington.

VA 22202-4302. and to the Office of Management and Budget. Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

November 1997

3. REPORTTYPE AND DATES COVERED

Technical Memorandum

5. FUNDING NUMBERS 4. TITLE AND SUBTITLE

An Impact-Location

Aircraft

Estimation Algorithm for Subsonic Uninhabited

6. AUTHOR(S)

Jeffrey E. Bauer and Edward Teets

7.PERFORMINGORGANIZATIONNAME(S) ANDADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9.SPONSORING/MONITORING AGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU 529 10 04 00 29 00

8, PERFORMING ORGANIZATION

REPORT NUMBER

H-2200

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/TM-97-206299

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 03

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

An impact-location estimation algorithm is being used at the NASA Dryden Flight Research Center to support

range safety for uninhabited aerial vehicle flight tests. The algorithm computes an impact location based on the

descent rate, mass, and altitude of the vehicle and current wind information. The predicted impact location is

continuously displayed on the range safety officer's moving map display so that the flightpath of the vehicle

can be routed to avoid ground assets if the flight must be terminated. The algorithm easily adapts to different

vehicle termination techniques and has been shown to be accurate to the extent required to support range safety

for subsonic uninhabited aerial vehicles. This paper describes how the algorithm functions, how the algorithm

is used at NASA Dryden, and how various termination techniques are handled by the algorithm. Other

approaches to predicting the impact location and the reasons why they were not selected for real-time

implementation are also discussed.

14. SUBJECTTERMS

Flight termination, Flight testing, Impact prediction, Range safety, Uninhabited

aircraft

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-2fl0-5500 Available from the NASA Center for AeroSpace Information, 800 Elkridge Landing Road,

Linthicum Heights, MD 21090; (301)621-0390

15. NUMBER OF PAGES

29

16. PRICE CODE

AO3

20. LIMITATION OF ABSTRACT

Unlimited

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI SId. Z3g-t8

298-'f02