Copyright ©1996, American Institute of Aeronautics and Astronautics, Inc.

AIAA Meeting Papers on Disc, July 1996A9635819, AIAA Paper 96-3491

The simulation technology for preflight evaluation of tactical missiles

Kent WangStandard Missile Co., McLean, VA

AIAA Flight Simulation Technologies Conference, San Diego, CA, July 29-31, 1996

We trace the conception and development of the preflight simulation technology for a tactical surface-to-airmissile. The applied concept of the flight simulation technology for the missile system is described, and the roundlevel of preflight performance for a tactical missile is also given. (Author)

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THE SIMULATION TECHNOLOGY FOR PREFLIGHT EVALUATION OFTACTICAL MISSILE

Kent WangStandard Missile Company

Abstract

Although the defense industry shrank in recent years, the application of flightsimulation technology to conduct prejlight evaluation remains intense. Theflight guidance and control systems benefited greatly from engineeringdevelopment simulation that provided not only high-fidelity representations offlying vehicle guidance and control, but also complex environments consistingof avionics, aerodynamics, and full scenarios. As a result, the value ofcomprehensive flight simulation as a critical aid to preflight evaluation wasfirmly established. This paper traces the conception and development of thepreflight simulation technology for a tactical surface-to-air missile. Theapplied concept of the flight simulation technology for missile system isdescribed, and the round level of preflight performance for a tactical missile isalso given.

I. Introduction

Models and simulations constructed for thepreflight evaluation of a tactical guided missileare often the best technical representation,outside of accumulated testing, of systemcapabilities and -limitations. They range incomplexity from two-dimensional/three degreesof freedom to three-dimensional/six degrees offreedom trajectory simulations. The hardware-in-the-loop simulations are normally designed tooperate in real time; this allows use of missilehardware in the simulation. Typical hardwarethat may be inserted in these simulationsincludes the steering and roll autopilots,guidance computer and receiver. Digitaltrajectory simulations are generally slower thanreal time, or utilize simplified aerodynamics andautopilot models to operate faster than real time.

Simplified kinematic simulations are often usedin test planning to define launch parameters andpredict impact conditions for unguided testvehicles. Specialized simulations are used toexamine critical flight phases such as launchand booster-missile separation. More complex

kinematic simulations are needed to developdesired altitude-velocity-maneuver profilesneeded to test control system operation. Closedloop guidance simulations are used to selecttarget flight profiles which will exercise specificfeatures of the guidance system. As they maturealong with deployment and exercise of thesystem and when combined with the judgmentof the simulation analysts, they become powerfultools for prediction of system performance. Notonly can they be used to predict missile systembehavior in the test environment, they can alsobecome very realistic estimators of systemperformance.

It is a well-known fact that there exist manymodels, simulations, and evaluation facilitiesthat are in use on a day-to-day basis throughoutthe tactical missile community. The informationcontained and derived from these tools isvaluable throughout the life cycle of the tacticalmissile from concept formulation throughengineering development, and particularly forpre-flight evaluation.

n. Description of Typical Guided Missile

A typical guided missile configuration isassumed for the following discussion. Themissile is surface launched. The airframe is

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cylindrical, and is maneuvered by moving fourindependent cruciform-mounted control surfacesfor both roll and steering motion. These controlsurfaces are mounted at the tail of this typicalmissile. Pitch and yaw steering is accomplishedby moving one pair of tail surfaces in the samedirection. Since this missile normally flies inhigh attitude, both pairs of control surfaces maybe deflected to accomplish pitch or yaw steeringmotion. Roll motion is effected by deflectingopposing tails in opposite directions. One pair ofcontrol surfaces defines a control plane. At thesame time, maneuvers in each control plane aredeveloped in response to commandaccelerations. These maneuvers may be in anypolar direction normal to the longitudinal axisof the missile.

Typical flight consists of four phases: boost,midcourse, acquisition and terminal homing.During boost phase the missile may follow afixed course determined by the launchingsystem. For a two-stage missile, roll stabilizationwill occur after separation. Roll stabilization isthe correction of the missile in-flight rollattitude to place the missile in the properattitude, and also compensate for undesirablerotation about its longitudinal axis. Roll errormay be produced by the launching system or in-flight aerodynamic forces. A typical roll controlsystem contains roll-sensing free and rate gyrosand a rate servo.

Prior to launch, the missile receives informationabout the target from the fire control systemthrough an umbilical connector on the launcher.If the midcourse mode is not utilized, this willrepresent the only target information it receivesuntil homing guidance begins. If the midcoursemode is utilized, target information will beupdated for a later initiation of terminal homing.

At boost phase termination, the missile may flyin a midcourse mode against long-range targets.Here the fire control system tracks the target andmissile, and computes missile maneuvers tomaintain it on the desired course. The necessarymaneuvers are uplinked to the missile until it iscommanded into its terminal guidance mode.

Terminal guidance for this typical missile issemiactive homing. The missile guidance systemis initially looking toward a point where thetarget is predicted to be. During the homingphase an illuminating radar is mechanicallyslaved to a track radar which is following thetarget. The illuminator energy reflected by thetarget is received by an antenna in the missilenose which defines the angles used in homingguidance. This front antenna is part of agimballed null-seeking, angle-tracking systemtermed a seeker. In two dimensions, the seekerspace angle is the sum of the missile bodyelevation and the seeker angle with respect tothe body. When the true line of sight is rotatingat a constant rate, the seeker head must rotate atthe same rate to continue tracking.

Deviations between the seeker antenna lookangle and the missile-target geometrical line-of-sight are represented by antenna pointing errorsignals derived by processing of the reflectedilluminator energy received from the target. Areference antenna located aft receives directenergy from the illuminator. The reflectedsignal and the direct signal provide the dopplerfrequency that is measure of closing velocity.The closing velocity estimation is used to modifythe guidance gain. A measure of the angularline-of-sight rate, is obtained from the sum of thehead rate gyro output and the rate of change ofthe tracking error. Once the reflected targetsignal is properly located by the tracking system,the terminal homing phase begins.

The missile is directed toward a predicted pointwhich it and the target will reach at the sametime. After the initiation of the terminal homingphase the missile is flying a collision coursewith the target, and the missile-to-targetazimuth and elevation line-of-sight angles donot change significantly unless the target coursechanges. Deviation from this collision coursedue to target maneuvers or missile errors aredetected by the guidance system which thencorrects the flight path to establish a new course.Semiactive homing with proportional navigationguidance is utilized in the terminal phase offlight.

During the homing phase the measure of line-of-sight angular rate is used to form commandmaneuvers. If either the target or missile changecourse, an angular change between missile

heading and missile-target line-of-sight willresult. Target course change may be an evasivemaneuver while missile course change may bethe result of internal errors. The angular changeis sensed and transformed to control surfacedeflections to maneuver the missile onto a newcourse. The command maneuvers are generatedin the guidance computer which also appliedfiltering to the noisy angular rate estimate thatincludes target amplitude and angularscintillation effects as well as receiver noise.The filtered signal is then multiplied by a gainproportional to the closing rate to generate thecommand acceleration. The terminal homingphase continues until missile and target arrive atthe same spatial point at which warheaddetonation terminates the flight.

in. Types of Flight Test

The effect of a propulsion unit on an existinglaunching system is evaluated by the flight oflaunching system. Evaluation of the flightperformance and aerodynamic history of apropulsion unit is tested by the flight of apropulsion test. This type of test is unguided, butmay have roll and pitch control autopilots toreduce dispersion effects and maintain stabilityfor marginally stable airframe flight conditions.Test data gathered in addition to propulsioncharacteristics include evaluation of dragmeasurements and booster-missile separationcharacteristics. Critical items for evaluationinclude launcher clearance, fire control systemcomponent clearance, and induced missile tip-off rates. Induced rates can be severe if thelaunching system is non-stabilized and subjectto rolling and pitching motion as aboard a navalvessel.

Cqntrol system performance is evaluated byflight of a control test which is unguided butcontains a maneuver program. This programproduces roll, pitch and yaw steering maneuverssingly and in combination. Response andstability characteristics of the roll and steeringautopilots are compared to- the predicted values.Additionally, there is a fallout of aerodynamicand propulsion data which can be compared topredictions.

Round level performance is evaluated in a flighttest round which utilizes a drone target flying a

predetermined velocity-altitude-range profile.Guidance system acquisition, response, stabilityand accuracy characteristics are compared topredictions over the flight profile. The premiertest criterion is the accuracy of placing themissile within a specified distance of the targetat the point of closest approach. In the event awarhead is used, the operation of the targetdetecting device as well as the damage inflictedon the target is the ultimate success/failcriterion.

The performance of variouscomponents/operations in the missile ismonitored by the use of telemetry to transmitmissile in-flight data to the ground. A commontelemeter unit used is a single radio frequency,pulse amplitude modulation (pam) or pulse codemodulation (pcm) system operating in the E-band frequency range. The telemeter unit maybe placed in the warhead location for test flights,and contains its own battery. It normallyradiates from two diametrically opposedantennas to provide complete pattern coverage.Supplemental instrumentation may be add to themissile to monitor specific phenomena such asvibration or thermal environment. Telemetryfunctions normally include instrument outputsand guidance and control system variables.Additional coverage of flights test is provided byradar and camera tracking.

IV. Description of Simulations

The simulation of the dynamic performance of amissile in six degrees of freedom requires arigorous three-dimensional aerodynamicrepresentation. This means completeaerodynamic data for the booster/missilecombination, missile during separation, andmissile alone. The aerodynamic data is requiredthrough the operating speed regime, for the fullrange of angle of attack, at all aerodynamic rollorientations, and for the range of steering androll control surface deflections. These data areobtained via aerodynamic testing usingtechniques of interpolation and extrapolation.

These aerodynamic data are analyzed to providea tabular three-dimensional aerodynamicdescription of the airframe for representation inthe 6 DOF trajectory simulation. Datarepresented include normal and side force

coefficients, pitching, yawing and rollingmoment coefficients, zero-lift drag coefficientfor all flight phases and the increment due tocontrol surface deflection and the angle ofattack. These coefficients are non-linearfunction of Mach number, angle of attack,control surface deflection in roll and steering.Mechanization of the kinematic equations in thesix degrees of freedom simulation isaccomplished by use of direction cosines.

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Figure 1 Simplified Stock Diagram of a Typical Homing Guidance Missile

Figure 1 presents a simplified block diagram ofthe simulation of a typical homing guidancemissile. Target and missile positions arecompared, and produce a tracking error. Therate of change of tracking error is summed withthe rate at which the seeker antenna must moveto point at the target; this summationapproximates the target line of sight angularrate. The angular rate and the missile-targetclosing rate estimate are inputs to the guidancecomputer which produces a command maneuverto the steering autopilot. The autopilot developssteering control surface deflection, and thisresults in achieved acceleration as well as drag.The loop is closed via the necessary geometricalsolutions and coordinate transformations.

In a digital simulation the computer isprogrammed to solve the algebraic equationsand numerically integrate the differentialequations with a computing interval smallenough to ensure an accurate solution. A typical

six degrees of freedom hybrid simulationoperates in real time and may contain a receiver,signal processor, guidance computer or autopilotwhich are either breadboard versions or flighthardware which is to be tested prior to use in amissile for the preflight test.

In summary, the six degrees of freedomsimulation is used to exercise a missile throughits complete sequence of flight modes and overits complete regime of flight conditions.Additional simulations utilized in preflightanalyses include commercial program whichanalyze transfer functions by generating rootlocus, frequency response or time response plots.These tools are also used in the analysis of thevarious control loops in the missile. They areused in conjunction with static aerodynamicsand dynamic condition. This allows the steeringautopilot designer to analyze the stability of thissystem at critical portions of the flight envelope,and also to determine the adequacy of theresponse characteristics to different types andmagnitudes of inputs. It also allows examinationof the effect of critical error sources on thedesign at static conditions.

V. Aerodynamic Test Analyses

The main objective in the test flight of apropulsion unit is to verify that the itsperformance is as predicted. Additionalobjectives include: (1) evaluation of booster-missile separation characteristics; (2)verification and checking of wind tunnelmeasurements of drag; and (3) validation of theflight dynamics simulation.

Some specific objectives are also set based onthe type of data desired. High altitude/maximumspeed data may be needed to verify the thermalenvironment. Burnout of the sustainerpropulsion unit some time before impact will bedesired to obtain power-off drag data. Smallangle of attack throughout flight will be neededto prevent induced drag from influencing thezero-lift drag measurements.

The test of a launching system is highlydependent on its environment. If the missile isto be launched from a non-stabilized launcherwhich also serves as a storage container for themissile, it is necessary to gather test data

describing the forces affecting the launchingsystem. If the launching system is mounted onthe deck of a fast and highly maneuverable ship,the test data must be introduced into the analysisas a time-varying input arid large samples oftypical data must be examined. Such an analysiswill require a three-dimensional, six degrees offreedom simulation to obtain the high degree ofaccuracy required. The simulation will solve theequations of motion of both missile andlaunching platform.

Launcher motion data will have been obtainedvia special instrumentation mounted on a testlauncher and recorded on magnetic tape forsimulation analyses. The data are normally inthe form of three linear accelerations and threeangular rates measured at several locations onthe ship with varying ship headings, ship speedand sea state. The instrumentation provides sixdegrees of freedom data at the test point. Thesedata must be transformed to define the motion ofthe ship at center of gravity location of themissile, and then input to the kinematicsimulation. Continued motion and velocity ofthe launching system components must also becomputed along with the position of criticalportions of the missile such as the controlsurface tips. These data will be referenced to aninertial reference system, and relative clearancebetween missile and these components will bedefined as the difference in inertial positions.The missile trajectory will be computed in twophases. During the first phase the missile will beaccelerating axially along the launch rail afterpropulsion unit ignition; here, the launcher isimposing forces normal to missile axis due toship motion. In the second phase, after themissile leaves the rail, its flight path isinfluenced by the lateral accelerations andvelocities imparted at the instant of release fromthe rail.

Clearance is established by relating missile andlaunching/fire control system positions atcommon points in time. Different views of themissile motion may be required to define criticalclearance. Initial analyses may assume thelauncher and missile are rigid bodies. However,launcher elastic measurements will yield theincremental effect of elastic deflections whichshould be included. Elastic deformation for thisproblem is a function of sea state, inducedvibration due to ship's speed, launcher elevation

and azimuth angles on the ship deck, and thefree bending modes of the ship.

The nominal altitude-speed profile can beattained by setting the elevation launch angleand programming small maneuver if adequatestability is maintained by inclusion of a steeringautopilot. Once the nominal trajectory isdefined, error sources must be considered. Errorcontributors include launch angle deviation andthrust and control surface misalignments for anon-stabilized propulsion unit. If a marginallystable airframe requires the use of a steeringcontrol autopilot with a null commandmaneuver input, additional error sources appear.Typical accelerometer bias is 0.1 g, while typicalrate gyro bias is 1.5 degrees/second. Both ofthese biases represent three standard deviationsof a Gaussian distribution, and 3 a encompasses99.73 percent of all possible values.Evaluation of the propulsion unit performancerequires adequate flight information such asradar range data, telemeter propulsion unitchamber pressure, and telemeter flightaccelerometer data. Since test flights of the newpropulsion unit present differences fromsimulation models in thrust versus time history,atmospheric conditions, and wind at altitude, itis necessary to determine these during flight andanalyze the data in detail to verify thepredictions. Telemeter chamber pressure andaccelerometer histories are used to calculate aninitial thrust-time history which is the input to apost-flight trajectory simulation. The outputvelocity of this simulation is compared withradar range data, and the difference is used togenerate an incremental acceleration andincremental thrust. The simulated thrust-timehistory is then adjusted using this incrementalthrust until a thrust profile is obtained whichmatches the flight. For the booster stage atmoderate launch angle, a thrust profile produceda total impulse that was less than one-halfpercent different from the nominal value andwell within acceptable tolerance variations.

Booster-missile separation produces a severeenvironment for the control system. Verificationof the separation dynamic model is essential. Animportant factor affecting separation is thestability of the airframe during the separationprocess. Here, missile speed is high and altitudeis relatively low; this produces a high dynamic

pressure environment. The airframe will beassumed to be aerodynamically unstable at thishigh q condition. This instability will producean angle of attack and body angular rate whichthe autopilot must be capable of capturing from.Additionally, roll rates accrued during theuncontrolled boost phase will subject the rollcontrol autopilot to further exertion.Complications arise when the booster is close tothe missile because it changes the airflowaround the missile and, thus, modifies theaerodynamics. As the two separate there is acontinually varying modification of theaerodynamics as the separation distanceincrease. The rate at which separation occurs isimportant since it determines the instantaneousmissile dynamics at the instant the autopilotcontrol loops are activated. These and otherconsiderations require a six degrees of freedomsimulation for prefiight analysis.

Other factors must be also considered in theanalysis are: (1) initial' conditions for positioncomponents as well as translational and angularvelocities; (2) variations in ambient temperaturewhich affect the propulsion unit; (3) frictionwhile on the rail; (4) cross winds;. (5) gravity;and (6) aerodynamics.

VL Guidance and Control Systems Analyses

The purpose of a control system test is toconfirm the characteristics of the missileairframe and roll and steering autopilot in adynamic environment. The specificcharacteristics to be tested include: (1) rollsystem capture and longitudinal stability duringand after booster separation; (2) steering systemstability during small and large commandmaneuvers; (3) roll and steering autopilot timeresponse over the expected range of dynamicpressure; (4) verification of aerodynamic data.

These objectives are achieved with a maneuverprogram provided by simulation analyses. Thismay include shifting the spatial roll attitudereference to check roll autopilot response, orconfining a maneuver mainly to the verticalplane for range coverage aspects. The roll anglemay be rotated while the missile is at large angleof attack such as may be experienced duringhoming guidance. This is equivalent to varying

the sideslip angle while the missile is at angle ofattack.

The steering autopilot simulation model iscarefully verified by comparing the time it takesto reach 63 percent of steady state in response toa step command maneuver; this is called theautopilot time constant. Stability of the controlsystem is verified by the absence of unplannedoscillations during the maneuvers. Uncertaintiesmay be exist in the induced rolling moment dueto sideslip angle; this is partly due to thedifficulty in accurately setting model controlsurfaces. An estimate of the validity of the windtunnel data was obtained by rolling a typical testvehicle to the single plane attitude andexecuting a fixed maneuver initially up-left, andthen rotating the maneuver to the up-rightdirection. As the maneuver progressed, theinduced rolling moments were generated and themissile began to roll. The rolling motion wasdetected by the roll control autopilot whichgenerated roll control surface motion to counterthe induced moment. The roll control surfacedeflection is then proportional to the inducedrolling moment.

The test objective for a guidance system is todemonstrate its operability as a tactical weapon.Specific test objectives are to verify guidanceaccuracy against a particular target scenario;this may test certain features of the guidancedesign such as acquisition and trackingcapability of the receiver against a small target;or the ability of guidance computer to responseto target evasive maneuvers. Due to the manyfacets of the guidance accuracy problem,extensive simulation analyses are required in theprefiight effort.

One of the most difficult problem is toaccurately simulate the target return signalsensed by the missile receiver and process this tothe form used as the input to the guidancecomputer. This signal processed by the guidancecomputer of the typical semiactive homingmissile will contain errors due to receiver,fading and glint noise. Receiver noise is afunction of target radar cross-section,illuminator-target range, and the missile-targetrange. Target fading noise is also termedamplitude noise, and is due to the variation inthe envelope of the target return signal producedby small scale target motion. Glint noise

represents the apparent motion of the control ofthe target-reflected energy sensed by the missile.Models for this random perturbation in thetracking error must be developed prior to pre-flight analyses. These nury differ from thoseused in the design study because targets forpreflight test are often small drones. Since all ofthe noise inputs are random time-varyingperturbations, this indicates the pre-flightanalyses must be of a stochastic nature.

Another source of error which will affect guidedflight is the missile internal variations such asbias and gain values associated with allelectronic equipment. These are termedhardware tolerances, and result from allowableinstrument and associated electronics variationsdefined in manufacturing specifications.One of the biases which most affects the typicalhoming missile using proportional navigation isthat due to the seeker head rate gyro. This biascan significantly perturb the flight oath byaltering the rate at which the apparent targetline-of-sight angle is changing. This effect mustbe considered in flight test planning.

A representative sources of instrument andelectronics tolerances which will affect theguidance accuracy of the typical missileincludes: (1) the seeker rate gyro bias and gaintolerances; (2) bias in the guidance computer;(3) errors in the steering autopilot (due mainlyto accelerometer and rate gyro errors); (4) one inthe roll loop; (5) fire control system; and, (6)variation in the propulsion unit total impulse.These tolerances are modeled as Gaussian-distributed with a mean equal to the nominalparameter value, and a standard deviationspecified for the manufacturing process. Thestandard deviation includes environmentalvariations which will perturb the hardwarecomponent.

The existence of these sources of randomvariations, noise and tolerances, plus thepresence of non-linearities such as a limit on thecommand acceleration requires a statisticaldetermination of guidance accuracy. Themethod used for the typical missile is to performa Monte Carlo analysis where the error sourcesare allowed to vary randomly within theirtolerance limits. Each combination may bethought of as an independent missile. A largenumber of these randomly selected missiles will

then be flown by simulation, and the statistics ofthe miss distance results will be determined. Nocorrelation is assumed between individualtolerances on a given simulated missile orbetween tolerances on successive missiles. Atypical Monte Carlo set consists of 50 flights,each representing an individual missile pluslaunching system conditions. Each flight alsoincludes the three random noise sources.

In addition to these tolerances, errors which alsoaffect guidance are those caused by therefractory properties of the radome which housesthe seeker antenna. This produces an errorbetween the seeker-target line-of-sight and thetrue missile-target line-of-sight, and thusgenerates an error in the seeker head rate gyroterm into the guidance computer. This error is afunction of many variables including radomethickness, radar frequency and polarization, andthe incidence angle of the signal on the radome.Manufacturing tolerances will alter the radomewall thickness from unit to unit. Therefore, theguidance accuracy statistical analyses alsoinclude variations in the radome-induced error.

Another variable which may affect guidanceaccuracy is the time at which homing guidanceis initiated. This depends on the time requiredfor the signal processor to search the videofrequency band for the target doppler, locate it,and lock on it. This requires accurate pointing ofthe seeker antenna during the search process,and the ability of the receiver to recognize thetarget signal above the noise level. Most MonteCarlo analyses assume that target acquisitionoccurs at a fixed time for a specific target; foranalyses where this time is critical, a hardwarereceiver and signal processor will be introducedinto the simulation. These hardware additionsare also required for the evaluation of guidanceaccuracy in the face of target electronic counter-measures.

VII. Ordnance Test Analyses

Factors which must be considered in theordnance test analysis are: (1) target size andtoughness; (2) target spatial properties; (3) fuzesignal return; (4) target hit point; (5) deliverymethod (exclusive hit-to-kill; exclusive stand-offmode or miss/graze/direct hit); (6) closing

velocity vectors; (7) aspect of closing conditions;and (8) miss distance.

Ordnance test analyses program simulates theaction of a target detecting^ device and warheadin approximately the last 100 milliseconds of amissile-target encounter. This program usedprimarily for computing probability of kill andfor detailed analysis of a method for solvingfuse-warhead coordination problems. Subsystemsimulated includes fuse and warhead, evaluateskill/damage probabilities due to lethal agent (rodor fragment) impact, blast overpressure, andcollision of missile and target.

The intercept data from the flight test provideskey missile and target variables at point ofclosest approach of the missile to the target or-tothe splash point, if the missile intersects thesurface. The point of closest approach is definedas the point where the missile to target closingrate equals zero. Along with the intercept data,also included the time history data can beprovided on the interface. The fallacy of thedistance domain is that it assumes a lineartrajectory to the intercept point; and thuslinearly backs up from the intercept point to thefuzing point. The fuze program contains thetime delay and burst point calculations. Thewarhead program uses the burstpoint data fromthe fuze program along with the warheadcharacterization data and target vulnerabilitymodels to calculate the probability of damagingthe target.

Vlli. Round Level Performance

The round level simulation is used to assesskinematic and guidance system performance forflight test predictions and also to verify systemdesign requirements. The simulation providesdetailed performance of the guidance andcontrol system, subsystem performancecharacteristics, and generates endgamegeometry information.

Round level simulation has features that make itflexible to meet a wide range of studyrequirements, e.g., it has the option of being runwith full three-dimensional aerodynamics ortrim aerodynamics. Specific subsystems such asthe receiver, seeker, autopilot, and inertialinstrument unit can be represented as simple or

complex models. Subsystem tolerances andnoise sources can be individually activated orredefined. The output of simulation includesmiss distance data, time histories of generalflight parameters and subsystem characteristics,and data in a format suitable for input to theprobability of damage program to determinesystem lethality. The simulation is used forflight test predictions, production ofperformance contours, and analysis of proposedfuture missile designs that use the basic missilestructure.

Subsystems of round level simulation provide adetailed depiction of all the major componentsof the missile. These include the radome, seekerantenna pattern, high-fidelity seeker, front FFTreceiver, digital signal processing algorithms,guidance computers, inertial instrument unit,autopilot, propulsion systems, control actuatorsystems, and aerodynamics for the boost,separation, and upper stage portions of flight.

Other subsystems of generic missile includingplatform, post-boost rocket motor, autopilot withcoupled aerodynamics, noise, sensitivities to gand angle rate, misalignment of axes foraccelerometer and gyro. Additional subsystemwill also affect round level flight test is theguidance system such as ship uplink model,guidance filter and law for midcourse, terminalbased on proportional navigation, modifiedbiases-guidance laws developed to showfeasibility of sidemount seeker.

Environmental factors include diffuse multipath,random ship motion, a statistical model of windvariation, clutter, receiver thermal noise, targetfading and glint models, variation of thrust dueto temperature, and radome boresight errors dueto in-flight heating.

Tolerances include statistical tolerance modelsfor the following: instrument biases,aerodynamic uncertainly, actuators, IRUalignment and drift errors, missile weight,missile center of gravity offset, vortex shedding,seeker head rate gyro bias, gain, andacceleration sensitivity, autopilot bias, andtracking errors.

In addition to the above, the simulation containsmodels of threat trajectories, maneuveringtargets, canister flyout, an interface to the

endgame lethality computer program, a splineleast squares fit algorithm for target trajectoryreconstruction, and logic for deterministiccalculation of target radar cross section andassociated geometric models of specific targets.

IX. Safety and Environment Consideration

Test vehicles are flown to test the effect of amissile performance on a launching system anda propulsion unit and also to confirm thecharacteristics of the missile airframe and rolland steering autopilots in a dynamicenvironment. Prior to a flight test from alaunching system, extensive analyses will havebeen conducted to insure that the design isadequate particularly in the context of safetyduring the launch phase. The concern is directphysical clearance between missile and launcheras it travels the launch rail, and missile and firecontrol system as it continues its flight afterleaving the rail.One of the requirements for testing tacticalmissile at any U.S. range is to provide positiveassurance against impact in an area occupied bynon-test personnel. Acceptable proof of safeimpact requires trajectory analyses of the missilein its most unfavorable failure mode, andseparate analyses of missile debris after breakupby the command-destruct system. Trajectorysimulation is used to provide theses data.

Null autopilot trajectories provided by a digitalsimulation are used to define forward, back andside range impact boundaries of the completemissile for range safety. These impactboundaries arc normally obtained by assumingrealistic failures in the electronics during flight.

Flight safety rules at test ranges require that aflight termination system be incorporated in themissile to allow command destruction if thevehicle exhibits anomalous behavior. Thewarhead itself may be used, or if a telemetry unitis in its place an explosive charge is required.Detonation of this charge will destroy airframestructural integrity, and cause the missile tobreak up in flight. The impact range of variousdebris elements is required. The method ofanalysis is to postulate a malfunction that causesthe missile to veer off course at its maximumlateral acceleration, and to assume major debriselements.

X. Conclusion

The simulation methods described for thedevelopment of a typical tactical missile providea cost-effective means of predicting flight resultswith information available prior to the test.Excellent agreement with flight data is possiblegiven that all test conditions are known and careis taken in their simulation. Aerodynamic datamay be verified or filled in by careful planningof flight maneuvers. Comparison with testresults also allows the simulation models to beupdated or corrected. New threat scenarios orflight environments may be easily investigatedor identified. In summary, without effective anddetailed simulations for pre-flight planning thegreat strides taken in providing sophisticatedtactical missile would not be possible.

References:

1. Albert, I, "Miss Distance Analysis forCommand Guided Missiles" Journal ofGuidance, Control and Dynamics, AmericanInstitute of Aeronautics and Astronautics,Vol.11, Nov.-Dec. 1988, pp. 481-487.

2. Eichblatt, E.J., Jr. "Test and Evaluation ofthe Tactical Missile", Vol. 119, Progress inAstronautics and Aeronautics series, AmericanInstitute of Aeronautics and Astronautics, NewYork, 1989.

3. Garnell, P., and East, D.J., "Guided WeaponControl System", Pergamon, Oxford, 1977.

4. Hemsch, M.J. and Nelson, J.M. "TacticalMissile Aerodynamics", Vol. 104, Progress inAstronautics and Aeronautics series, AmericanInstitute of Aeronautics and Astronautics, NewYork, 1986.

5. Nesline, F.W., and Zarchan, P., " MissDistance Dynamics in Homing Missiles",Proceeding of American Institute of Aeronauticsand Astronautics Guidance and ControlConference, American Institute of Aeronauticsand Astronautics, New York, Aug. 1984.

6. Zarchan, Paul "Tactical and StrategicMissile Guidance", Vol. 124, Progress inAstronautics and Aeronautics series, AIAA,New York, 1990.