Evaluation of Guidance and Control System of High

Speed Flight Demonstrator Phase II

Tetsujiro Ninomiya∗, Hirokazu Suzuki†, and Taro Tsukamoto†

Japan Aerospace Exploration Agency, Mitaka, Tokyo, 182-8522, Japan

This paper describes the results of the preflight evaluation of the guidance and controlsystems of the High Speed Flight Demonstrator Phase II (HSFD-II). Since the vehicledoes not have a nominal trajectory because of its launch method, it is not appropriate toevaluate this system by the root sum square method. The present evaluation followed twoapproaches, using one standard and one custom designed analysis. The former comprised aMonte-Carlo simulation, single error analysis, sensitivity analysis, and linear analysis. Thelatter used emergency separation, release oscillation, and GPS receiver error. Although insome cases the Monte-Carlo simulation in the standard analysis showed a failure to satisfymission requirements, detailed analysis indicates that the system would guide the vehiclethrough a successful flight experiment. This evaluation establishes that the system satisfiesall mission requirements. The actual flight experiment was carried out and the guidanceand control systems worked quite well. Based on the results of the preflight evaluation andthe results of the flight, this paper concludes that the guidance and control systems of thevehicle was properly designed.

Nomenclature

EAS Equivalent Air SpeedGC Guidance, and ControlGNC Guidance, Navigation, and ControlGPS Global Positioning System

I. Introduction

Japan Aerospace Exploration Agency (JAXA) has studied for the future space transportation. OREX(Orbital Reentry Experiment) was carried out in 1994 to investigate the aerodynamic heating at the reentryto the Earth, HYFLEX (Hypersonic Flight Experiment) was implemented in 1996 to demonstrate the GNCsystems in the hypersonic region, and ALFLEX (Automatic Landing Flight Experiment)1 was also put intopractice in 1996 to establish the autonomous landing technology. The objective of these experiments were todevelop technologies for HOPE-X (H-II Orbiting Plane Experimental), and HSFD programs were planned toobtain the fundamental knowledge for future space transportation systems.2 HSFD program consists of twophases; HSFD-I aims to inspect the environmental conditions of landing site, and HSFD-II is designed toobtain the aerodynamic data around the transonic region. This paper describes the evaluation of guidanceand control systems of HSFD-II.

HSFD-II campaign was curried out from March to July 2003 under the cooperations of NAL, NASDA,and CNES at Esrange experiment site in Sweden (NAL and NASDA were merged into JAXA in October2003).

The purpose of HSFD-II is to obtain the aerodynamic data of a winged re-entry vehicle around thetransonic region. Some flight experiments3 and wind tunnel tests show that the base pressure correction is

∗Scientist, Institute of Space Technology and Aeronautics, 7-44-1 JindaijiHigashi-machi Chofu-City 182-8522 Tokyo Japan†Associate Senior Researcher, Institute of Space Technology and Aeronautics, 7-44-1 JindaijiHigashi-machi Chofu-City 182-

8522 Tokyo Japan, Member AIAA

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AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies AIAA 2005-3274

Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

2H FD

Ascent ( 2.5 3 h )

Launch

Launch Preparation

Release ( Altitude 30km )

Data Acquisition

Deployment of Parachute

Deployment of Airbags

Touch Down

Ceiling ( 6.5 h ) Acceleration

Pre-Flight Check

DGPS Ground Station

Balloon Launch Site

Stratospheric Balloon

Telemeter / Command Antenas

Weather Station

GPS Satellites

Figure 1. Mission Profile of HSFD II

difficult for the space plane shape body because of the interference of model holding structure. Especially,the wind tunnel tests show that the dispersion of the results were twice as large as that was expected.This large error requires the large margin for the controllers. Moreover, since high speed wind tunnel testdemands smaller model, this results in the larger errors in measurement of hinge moment, and the largererrors requires larger margin for the actuator models. These threaten the feasibility of the vehicle.

High altitude stratospheric balloon was selected as the launch method for HSFD-II based on the compar-ison with some other launch methods.4 Figure 1 shows the outline of the experiment. The vehicle is similarto 25% size of HOPE-X vehicle and it is delivered to sufficient altitude hung by a stratospheric balloon. Thenit is released and fell down to acquire the objective velocity, and it obtains the aerodynamic data aroundthe transonic region. This experiment method is quite unique and there are few similar experiment before.At the same time, the development of this type of experiment has a drawback in the same way as the spacesystems, i.e., it is impossible to extend the flight envelop that is common method in the development of theairplanes. Accordingly, it is important to exhaustively evaluate the GNC systems beforehand by computersimulations.

This paper describes the evaluation method, and the result of actual flight experiment is shown.

II. Evaluation of GC systems

Since the launch method is stratospheric balloon, it is impossible to control the horizontal position atthe release of the vehicle. Moreover, it does not have the engine so that the gliding capacity is severelylimited. To settle this problem, some recovery points are set and it is required that the vehicle can glide toat least one of them wherever in the experimental area it is released. These facts result in that there are nonominal trajectory of this experiment and this directly means that it is insufficient to evaluate GNC systemsby single error analysis which is often used in the former experiments. Therefor Monte-Carlo simulation(MC) is applied as a principal evaluation method.

In this section, the overview of the guidance and control system of HSFD-II is described. Then standardand custom designed evaluation method are mentioned. The former is a part of guidance and control designiteration and it consists of linear analysis, Single error and sensitivity analysis, MC. The latter consists ofrelease altitude analysis, which are necessary because of the unique procedure of this experiment, oscillation

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at the separation, and GPS malfunction analysis. The success of the mission is judged by the criterionsummarized in Table 1. Required specifications are the criterion for MC and Surveillance items are additionalitems for this evaluation for the following reason.

Prohibited area intrusion Around the Esrange area, some dense population area is recognized as flightprohibited area. Although there is no rule to forbid all flight in this area, we have decided to check thepossibility of intrusion.

Excessive β Since our aerodynamic model is limited with respect to β in the range from −5◦ to 5◦, thereliability of the simulation gets lower under excessive β condition.

Excessive β during measurement Smaller β is desirable from the standpoint of data acquisition.

Table 1. Criterion of mission success

item success condition

Divergence The angle of attack keeps the range of −10◦ to 30◦ dur-ing whole flight

Data acquisition phase skip Mach number hold phase sustains more than 1s.

Recovery failure The point of parachute deployment is included in therecovery cone(see Figure 2).

Excessive EAS EAS at parachute deployment is less than 103m/s.Maximum dynamic pressure Maximum dynamic pressure is less than 15.68kPa.Max/min load factor Load factor keeps the range of −1.0g ∼ +3.5g.

Insufficient data acquisitionDuring Mach number hold phase, the velocity holds theobjective Mach number ±0.03 and sweep the angle ofattack from 10◦ to 2.5◦ with |α̇| ≤ 2deg/s.

Prohibited area Intrusion The vehicle keeps out of prohibited area.

Excessive βThe slip angle is within −5.0◦ and 5.0◦ during wholeflight.

Excessive β during measurement The slip angle is within −2.0◦ to 2.0◦ during Mach num-ber hold phase.

Req

uire

dsp

ecifi

cati

onC

onfir

mat

ion

A. Guidance and Control systems

H=1,500m

H=1,067m

H=0m

Recovery Area

(Di. 3km)

Recovery Cone

Recovery Cone

Figure 2. Recovery Cone

The guidance and control systems are summarizedin Reference.5 The flight of HSFD-II consists of6 phases as is shown in Figure 3. After releasefrom from the balloon, the first phase is accelera-tion phase, in which the vehicle keeps its pitch an-gle to be −80◦ and descends to accelerate. Thenit prepares for the aerodynamic data acquisition intrajectory insertion phase, and it obtains the datain Mach number hold phase. Then it gets nose up indeceleration phase, and then here comes the returnphase where it flies to the recovery point. The fi-nal phase is recovery phase and the recovery systemworks sequentially in this phase. This final phase isout of our concern in this paper.

The vehicle is model with 144 errors. Measured errors, such as mass, are modeled as uniform distributionand Estimation errors, such as aerodynamic coefficients, are modeled as normal distribution.

B. Standard analysis

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(1) Acceleration phase

Accelerates with minimum pitch angle(-80deg)

(2) Trajectory insertion phase

Pulls nose up to maximum angle of attack (13 deg)

(3) Mach number hold phase

Holds Mach number

Sweeps angle of attack

(4) Deceleration phase

Decelerates with nose up under maximum load factor

(5) Return phase

Selects recovery point

Turn if necessary

Flys along with reference trajectory of -20 deg

flight path angle

(6) Recovery phase

Figure 3. Mission sequence

Standard analysis comprised ofthree methods. As is describedbefore, the controller gains aretuned through MC. Single errorand sensitivity analysis respec-tively shows the sensitivity toa 3σ size error and the limit ofcontrol capability. However, itis important to analyze the sys-tem based on the linear controltheory, so that the linear anal-ysis is implemented.

The objective Mach num-bers are set at 1.2, 1.05, 0.8.The GC systems are evaluatedfor each Mach number.

1. Monte-Carlo Simulation

In MC, combinations of ran-dom errors for all error sourcesare assumed and simulationsare carried out. For each objec-tive Mach numbers, 1000 casesare executed and the results areexamined. Our goal is to re-duce the ratio of fatal failure,which might result in the vehicle destruction, to less than 0.3%.

The results of MC is shown in Table 2. This table is shown in priority order from the top, and failurecase is not repeatedly counted in required specification section.

Table 2. Results of MC

item 1.20 1.05 0.80Success 985 986 986Failure 15 14 14

RequiredSpecification

Divergence 0 0 0Data acquisition phase skip 0 0 0

Recovery failure 2 1 1Excessive EAS 9 10 5

Maximum dynamic pressure 0 0 0Max/min load factor 4 2 7

Insufficient data acquisition 0 1 1

ConfirmationIntrusion 2 0 1

Excessive β 7 2 12Excessive β during measurement 11 18 1

All failure cases are precisely investigated and the results are summarized as follows:

• 4 cases of recovery failure can be avoided by operation.

• 24 cases of Excessive EAS are very slight violation, and it is judged that the destruction risk of thesecases are small.

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• 3 cases of Mach 1.2 out of 13 cases of Max/min load factor are very serious violation of the criterion,and these case might result in vehicle destruction.

From the mentioned above, the frequency of mission failure is 39/30000, and 3/30000 are serious cases ofpossible vehicle damage. This evaluation indicates that the GC systems have required performance.

2. Single error and sensitivity analysis

We discuss both the single error analysis and sensitivity analysis in this section because of their similarity.In ALFLEX and HYFLEX project, single error analysis performed a principal evaluation.

Now, we discuss the outline of these method. Each error source is independent and nominal trajectorywith no error is assumed. In this case, the separation point is set at the center of experimental area. Thenone error is chosen and its value is set to be +3σ. Once simulation for this case is executed, the differenceof each evaluation item from that of nominal case is calculated. Same manipulation is done for −3σ errorand the worse case is hold. This process is done for all error sources and the root sum square (RSS) isadded to the nominal value. If this value satisfies the criteria, we can see that the GC systems have requiredperformance.

In the sensitivity analysis, error value of nominal distribution errors are extended to ±6σ and ±9σ, andthis is the only difference from single error analysis.

Required qualification for GC systems is no violation for single error analysis. The aim of sensitivityanalysis is to obtain useful information, so that we do not have any strict criteria.

In this analysis, the separation point is assumed to be the center of ZONE B area.The results of this analysis is shown in table 3. These results show that the GC systems have required

performance for single error analysis. The sensitivity analysis indicates that 6σ size errors never cause fatalerror except for Cmα+ error case.

Table 3. Violation cases of single error analysis and sensitivity analysis

Item 1.2 1.05 0.83σ 6σ 9σ 3σ 6σ 9σ 3σ 6σ 9σ

RequiredSpecification

Divergence 0 0 5 0 0 5 0 0 1Data acquisition phase skip 0 0 0 0 0 0 0 0 1

Recovery failure 0 1 1 0 1 1 0 1 1Excessive EAS 0 0 1 0 0 1 0 0 1

Maximum dynamic pressure 0 0 0 0 0 0 0 0 0Max/min load factor 0 0 1 0 3 1 0 1 1

Insufficient data acquisition 0 0 0 0 0 0 0 2 4

ConfirmationIntrusion 0 0 0 0 0 0 0 0 0

Excessive β 0 0 0 0 0 0 0 0 0Excessive β during measurement 0 0 0 0 1 2 0 0 1

3. Linear analysis

In this analysis, the nominal trajectory is assumed in the same way as the previous section. The vehicledynamics is linearized at the points on this trajectory in every one second. The stability and stability marginsare evaluated at each point. The criterion of this analysis is that the gain margin is more than 6dB and thedelay margin is more than 0.1s. This analysis is carried out for each target Mach number case.

For the longitudinal motion (Figure 4), the vehicle is inherently unstable without control. Stabilizationand the tracking to the pitch command is performed by elevators. The gain margins mostly hold morethan 5dB and the delay margins hold more than 0.1s. Exceptional points are the cases around 45s afterseparation. The angle of attack of these cases are almost maximum value and this fact causes the insufficientstability. However this condition lasts within one second.

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0 50 100 150 200 250 300 350-0.2

0

0.2

0.4

0.6

0.8

1

0 50 100 150 200 250 300 3500

2

4

6

8

10

12

14

16

18

20

Del

ay M

argin

[se

c]

Time [s]

Gai

n M

arg

in [

dB

]

Time [s]

Figure 4. Linear analysis : longitudinal motion for Mach 1.2 case

0 50 100 150 200 250 300 3500

2

4

6

8

10

12

14

16

18

20

0 50 100 150 200 250 300 350-0.2

0

0.2

0.4

0.6

0.8

1

Gai

n M

arg

in [

dB

]

Time [s]

Del

ay M

argin

[se

c]

Time [s]

aileronladder

aileronladder

Figure 5. Linear analysis : lateral and directional motion for Mach1.2 case

For the lateral and directional motion(Figure 5), the vehicle has the sufficientinherent stability except for the very lowspeed case (about 15s since separation)and low angle attack in transonic regioncase. In this exceptional region, the gainmargin of aileron and delay margin ofladder do not satisfy the criterion, how-ever, this violation continues only for 6seconds.

Summarizing up this results, al-though there are some interval where thecriterion is not satisfied, the other anal-yses show that this fact does not effectthe actual flight. Only if there is errorwhich enforces this violation, there is illeffects on the flight. We judged that thispossibility is low enough.

C. Custom designed analysis

In this section, we describe some extraevaluation for the cases which are notmodeled in a general way. The purpose ofthis analysis is not to feedback the knowl-edge to the guidance and control design,but to estimate the uncertainness causedby the unique experiment method.

1. Release altitude analysis

0

5

10

15

20

19 20 21 22 23 24 25 26 27 28 29 30 31 32 33

Nu

mb

er o

f d

iver

gen

ce

Release altitude [km]

Mach 0.8

Mach 0.95

Mach 1.2

Mach 1.05

Figure 6. Release altitude analysis : Standardseparation cases

In this analysis, we consider the case in which some emer-gency requires an immediate separation, while the vehicleis hung from the balloon and is ascending. In the standardanalysis, the separation altitude is assumed as nominal al-titude plus error, however, this assumption does not coverall altitude from the ground to 30km. If the vehicle is sep-arated at uncovered area, the vehicle starts the emergencyrecovery process, but this is not always critical situation.Accordingly, the flight envelope can be extended, and thisis the reason why this analysis is motivated.

The method of this analysis is quite simple. Thereare two characteristic altitude. One is the lowest altitude(18820m) to accelerate vehicle to Mach 0.8, and the otheris the highest altitude (32850m) to accelerate the vehicleto Mach 1.2. Two cases are considered using these twoaltitude; standard separation case and low altitude sepa-ration case.

In the former case, the separation error is assumedto be uniform distribution and the separate altitude isassumed to uniformly distribute between the highest andlowest altitude. Then MC with this modification is applied, and the result is shown in Figure 6.

In the latter case, almost the same procedure is applied. The separation altitude is assumed to uniformlydistribute between the lowest altitude and 4000m, which is determined by parachute operation. The resultis shown in Figure 7.

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0

50

100

150

200

250

300

350

10 12 14 16 18864

Nu

ber

of

div

erg

ence

Release altitude [km]

Figure 7. Release altitude analysis : Emergencyseparation cases

In both case, 1000 cases MC is carried out. Fromthese results, we can see the followings. There is a dan-gerous area (25km to 26.5km) in the standard separationaltitude, and if the emergency happens under 18820m, itis recommended that the emergency recovery procedurestarts immediately.

2. Release oscillation analysis

As we mentioned above, this flight experiment is quiteunique in the point of its launch method. There was onlyone case which have been operated by CNES. The releaseoscillation errors in our model refers to this experiment,however this model is not sufficiently reliable.

Consequently, the release oscillation analysis is imple-mented. This is also important for the emergency case,in which we have no time to wait for the vehicle to stopthe oscillation.

The evaluation method is as follows. The release os-cillation model is originally modeled as uniform distribu-tion, and it is converted to normal distribution with sameaverage and ±3σ being the upper and lower limit of uniform distribution. Then the sensitivity analysis andMC are put into practice. This analysis is done only for Mach 0.8 case by request.

The results of sensitivity analysis show that ±9σ of yaw rate cases cause the divergence. From this, theseparation error of pitch and roll rate are set to ±9σ and that of yaw rate is set to ±6σ, and MC is executed.This MC results in 45 cases divergence out of 1000 cases. In these divergence cases, each of roll, pitch, yawrate is set to 3σ value and they are simulated again. This re-simulation results indicate that if we can controlthe yaw rate within the current limit (±1.0deg/s), the flight does not fail under the condition of ±6.0deg/s(corresponding to ±9σ) roll rate and ±3.0deg/s (corresponding to ±9σ) pitch rate.

Table 4. Release oscillation analysis

Item Mach 0.8 3σ roll rate 3σ pitch rate 3σ yaw rateSuccess 672 6 28 35Failure 328 39 17 10

RequiredSpecification

Divergence 45 37 9 0Data acquisition phase skip 0 0 0 0

Recovery failure 19 2 2 3Excessive EAS 14 0 0 0

Maximum dynamic pressure 0 0 0 0Max/min load factor 3 0 0 0

Insufficient data acquisition 1 0 0 0

ConfirmationIntrusion 0 0 0 0

Excessive β 42 4 5 0Excessive β during measurement 1 0 0 0

3. GPS malfunction analysis

In this analysis, we assumed the discontinuity of navigation data caused by malfunction on the GPS receiver.There are two possible cases of this phenomena. At the separation, GPS antenna is switched from one onthe gondola to the other on the vehicle, and this might cause the navigation data discontinuity. This case

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is modeled as Figure 8. The other case is large angle maneuver of the vehicle. In this case, the satellitesconstellation changes and this cause the navigation data discontinuity. This case is modeled as Figure 9.

-40-20

02040

0 100 200 300

-40-20

02040

0 100 200 300

-60-30

03060

0 100 200 300

-0.4-0.2

00.20.4

0 100 200 300

Dir

ectio

nal

erro

r [d

eg]

Time [s]

Sout

h-N

orth

erro

r [m

]E

ast-

Wes

ter

ror

[m]

Alti

tude

erro

r [m

]

Time [s]

Figure 8. GPS malfunction analysis : separation case

-2

0

2

-2

0

2

-4-2024

-0.4-0.2

00.20.4

0 100 200 300

0 100 200 300

0 100 200 300

0 100 200 300

Time [s]

Sout

h-N

orth

velo

city

err

or[m

/s]

Dir

ectio

nal

erro

r [d

eg]

Eas

t-W

est

velo

city

err

or[m

/s]

Alti

tude

vel

ocity

erro

r [m

/s]

Time [s]

Figure 9. GPS malfunction analysis : large maneuver case

These causes are not a breakdown of nav-igation system, and this make it difficult tohandle. The objective of this evaluation is toestimate the possibility of flight failure causedby the navigation data discontinuity. From theresults of standard analysis, three less stablecases are selected to forcus.

1. High angle of attack area in the earlypart of Mach number holding phase (An-gle of attack is more than 12 deg.)

2. Low angle of attack area in the late partof Mach number holding phase (Angle ofattack is less than 3 deg.)

3. The area around the timing of assumedrecovery point change in the returnphase.

Concretely, one of two kind of navigation datadiscontinuity is added at the timing of upper1 to 3 cases, and five error sources which aremost sensitive to the flight failure are added.

The result is that there is only one violationout of 1440 cases. It is violation of maximumload factor, but the maximum load factor is3.05g where the limit is 3.0g. From this, wecan conclude that navigation data discontinu-ity has quite limited effect on the flight safety.

III. Flight results

68.0

68.2

68.4

68.6

20.2 21.0 22.0

50km0

67.8

Longitude (deg)

La

titu

de

(d

eg

)

Esrange

Kiruna

Ascent

Launch

Release

HSFD Flight

HAC

Touch-Down

Figure 10. Footprint of the flight

HSFD Vehicle ( July, 1 )68 13.071' N 21 08.999' E

Riser Cover ( July, 1 )68 13.302' N 21 06.020' E

Drogue Gun Activation

Recovery Area #05

HSFD Trajectory( Estimated Actual )

Drogue GUn Activation

Error due to Inertial Navigation

HSFD Trajectory( Telemetry Down-Link )

Drogue Gun Slug (July, 9 )68 12.876' N 21 04.163' E

Parachute Door Sub-Panel ( July, 2 )68 12.896' N 21 03.819' E

Figure 11. Landing point

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Time after Release t (s)

Alt

itu

de

H (

km

)

Ma

ch

Nu

mb

er

M

An

gle

of

Att

ac

k

(d

eg

)

Sid

esli

p A

ng

le

(d

eg

)

Constant Mach Phase

Figure 12. Flight status during the data acquisition

The first flight experiment was carriedout on July 1st 2003. The target Machnumber was 0.8. Figure 10 to 12 showthe actual flight data. In this flight, thevehicle could not receive the GPS dataduring its ascent, and this condition holduntil the end of the flight. Accordingly,the vehicle flew only with inertial navi-gation system, and it is estimated that itflew to 2.7km south to the selected recov-ery point (see Figure 11) because of thenavigation error.

Figure 12 shows the flight status dur-ing the data acquisition. These figuresshow that all requirements (see Table1)are satisfied during this phase. Machnumber holds the range of 0.8 ± 0.03,angle of attack sweeps the range from10.0deg to 2.5deg with less than 2.0deg/sratio, and the slip angle holds within±2.0deg.

The maximum dynamic pressure and the maximum and minimum load factor during the whole flightwere 9.57kPa,+2.24g,−0.0g respectively. The point of parachute deployment under the inertial navigationdata was within the recovery cone, and the altitude and position errors were 11.0m and 29.8m. The EAS atthe recovery point was 92.7m/s. All of these results show that the all criterion were satisfied.

From this result, we conclude that the GC systems worked properly.

IV. Conclusion

We discussed the evaluation method for the GC systems of HSFD-II and showed the flight results. Theevaluation results lead us to the following conclusions. GC systems are evaluated in standard evaluation, andthey show that the GC systems have sufficient performance for the flight success. Some extra analysis whichare called customized evaluation are implemented. These analysis indicate important information which isuseful for the flight operation.

In addition, not only we have developed the evaluation tools but also we have used them iteratively totune the guidance and control gains, to estimate some particular conditions which are not modeled usually,to determine the operational conditions, and so on. This thorough use of the evaluation tools is the key tothe success of GC systems development.

References

1“Proceedings of the ALFLEX Symposium,” NAL SP-39t, NAL, Tokyo, Japan, Aug. 1998.2Yanagihara, M., Miyazawa, Y., and Taniguchi, H., “HOPE–X High Speed Flight Demonstration Program Phase

II,” Aiaa paper 2001–1805, NAL, April 2001.3HYFLEX Aerodynamic Characteristics Research Team, “Aerodynamic Characteristics of Hypersonic Flight

Experiment (HYFLEX) Vehicle,” NAL TR-1334, NAL, Tokyo, Japan, Dec. 1997, (Japanese).4Yanagihara, M., Miyazawa, Y., and Taniguchi, H., “Simulation Analysis of the HOPE–X Demonstrator,” Aiaa

paper 99–4875, NAL, Nov. 1999.5Tsukamoto, T., Suzuki, H., Ninomiya, T., and Nishizawa, T., “GUIDANCE AND CONTROL FOR THE HIGH

SPEED FLIGHT DEMONSTRATION PHASE II,” Aiaa paper 2004–4944, Aug. 2004.

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