ISSN 1349-1113JAXA-RR-04-035E

JAXA Research and Development Report

Comparative Force/Heat Flux Measurements between JAXAHypersonic Test Facilities Using Standard Model HB-2

(Part 1: 1.27 m Hypersonic Wind Tunnel Results)

Shigeru KUCHI-ISHI, Shigeya WATANABE, Shinji NAGAI,

Shoichi TSUDA, Tadao KOYAMA, Noriaki HIRABAYASHI,

Hideo SEKINE, and Koichi HOZUMI

March 2005

Japan Aerospace Exploration Agency

慣性速度情報を用いたADS横滑り角の補正 1

Comparative Force/Heat Flux Measurements between JAXAHypersonic Test Facilities Using Standard Model HB-2

(Part 1: 1.27 m Hypersonic Wind Tunnel Results)

Shigeru KUCHI-ISHI*1, Shigeya WATANABE*1, Shinji NAGAI*1,

Shoichi TSUDA*1, Tadao KOYAMA*1, Noriaki HIRABAYASHI*1,

Hideo SEKINE*2, and Koichi HOZUMI*2

HB-2形状標準模型を用いた JAXA極超音速風洞間力／加熱率対応風試（その 1： 1.27 m極超音速風洞測定結果）

口石　茂 *1、渡辺重哉 *1、永井伸治 *1、津田尚一 *1

小山忠勇 *1、平林則明 *1、関根英夫 *2、穂積弘一 *2

ABSTRACT

This report presents a detailed description and results of force and heat flux measurement tests conducted in the 1.27 m

Hypersonic Wind Tunnel (HWT) of the Japan Aerospace Exploration Agency (JAXA). The HB-2 standard hypersonic bal-

listic configuration was employed as a model. The force-measuring tests used a stainless steel model on a six-component

balance, and the heat flux measurement tests were made using a chromel model with a total of 28 coaxial thermocouples

press-fitted on the surface. A non-intrusive heat flux measurement was also made by infrared (IR) thermography using an

alternative nose part made of polyimide plastic. The tests were conducted at a nominal Mach number of 10, a stagnation

enthalpy of 1 MJ/kg, and stagnation pressures from 1 to 6 MPa. For both force and heat tests, good data repeatability was

confirmed. The heat transfer coefficient obtained from the IR thermography agreed well with that from the thermocouple

measurement. For the heat flux data, uncertainties associated with flow repeatability, the model’s streamwise location in

the test section, and the model’s alignment were quantified and examined. Also, the measurement error of both the force

and heat tests was evaluated. A conventional statistical approach which estimates the bias and random error components

was applied to the force test data, while a Monte Carlo approach was used to numerically estimate the uncertainty in the

data reduction process for the heat flux data. The present experiment was conducted as a series of comparison tests between

two hypersonic facilities in JAXA. Since the free-stream conditions and the corresponding experimental data were tabu-

lated in detail for each run, the present test data serve as a database not only for the evaluation of force and heat flux mea-

surement in HWT, but also for the validation of hypersonic computational fluid dynamics (CFD) codes.

Keywords: hypersonic wind tunnel, standard model, aerodynamic force, heat flux, uncertainty analysis

概　　　　　要

本報告は、JAXA 1.27 m極超音速風洞（HWT）において実施された、HB-2形状バリスティック標準模型を

用いた空気力/空力加熱率測定試験結果に関するものである。加熱率測定試験で用いられた模型はクロメル製

であり、計 28個の同軸熱電対が装着されている。模型頭部はベスペル製部品と交換し、赤外線カメラによる

非接触計測の結果と比較が可能となっている。試験は風洞澱み温度 1000 K、澱み圧力 1－ 6 MPaの範囲で実

＊ 1 Wind Tunnel Technology Center, Institute of Space Technology and Aeronautics（総合技術研究本部風洞技術開発センター）＊ 2 Foundation for Promotion of Japanese Aerospace Technology（航空宇宙技術振興財団）

NOMENCLATURE

A = Reference area, πD2/4

Ab = Model base area

B = Bias limit

CA = Zero-lift gross (total) axial force coefficient, Fx/q∞A

CAF = Zero-lift forebody axial force coefficient, CA－(p∞－

pb) Ab/A

CBm = Pitching-moment coefficient based on the balance cen-

ter, My/q∞AD

Cm = Pitching-moment coefficient based on the reference

point, CBm－ dCN/D

CN = Normal force coefficient, Fz/q∞A

c = Specific heat

D = Reference centerbody diameter (see Fig. 2)

d = distance from the balance center to the moment refer-

ence point

Fx = Axial aerodynamic force measured by the balance

Fz = Normal aerodynamic force measured by the balance

h = Heat transfer coefficient, q/(Taw－ Tw)

k = Thermal conductivity

L = Distance from the nozzle exit (see Table C2)

My = Pitching moment measured by the balance

M = Mach number

P = Precision limit, tS

Ppit = Pitot pressure

P0 = Tunnel stagnation pressure

p = Pressure

pb = Base pressure

q = Heat flux, dynamic pressure

R = Gas constant

Re = Free-stream Reynolds number based on centerbody

diameter

S = Precision index

T = Temperature

Taw = Adiabatic temperature

T0 = Tunnel stagnation temperature, initial temperature

t = Time, Student t value

t0 = Initial time

U = Velocity, total uncertainty, random variable

x = Axial distance

α = Angle of attack

φ = Roll angle

θ = Circumferential position

ρ = Density

γ = Specific heat ratio

Subscripts

w = Model surface (wall)

0 = Stagnation

∞ = Free-stream

1. INTRODUCTION

For the design and development of future hypersonic flight

vehicles, the prediction of aerothermal characteristics during the

atmospheric reentry is of importance. Due to limitations of con-

ventional ground-based experimental facilities, it is practically

impossible to produce flows of desired Mach number, Reynolds

number, and total enthalpy exactly the same as those in actual

flight conditions. As well, a numerical prediction applying com-

putational fluid dynamics (CFD) techniques to date is insufficient

in terms of reliability since it includes many uncertainties for

both numerical and physical aspects. Therefore, it is apparent that

the use of either experimental or numerical approach alone is

inadequate for a reliable flight prediction and it is the best way to

apply all of possible approaches and to evaluate the prediction

accuracy from a global point of view.

The Japan Aerospace Exploration Agency (JAXA) has two

large-scale hypersonic facilities called the 1.27 m Hypersonic

Wind Tunnel (HWT) and the High Enthalpy Shock Tunnel

(HIEST). These facilities are common in the sense that they cover

hypersonic speeds, but there are differences regarding the flow

JAXA Research and Development Report JAXA-RR-04-035E2

施された。力試験/熱試験ともに、気流再現性、模型射出位置、モデル取りつけ精度に関するデータのばらつ

きを定量的に評価した。データ再現性は良好で、HWTの良好な気流特性を示す結果となった。力試験につい

ては、過去実施された他風洞における試験データと比較し、良好な一致が確認された。また力試験、熱試験

共に不確かさ解析を実施した。力試験については従来用いられる統計的手法を適用し、得られた誤差幅は測

定で得られたデータのばらつきと比較して妥当であることが確認された。一方熱試験についてはモンテカル

ロ法による不確かさの数値的導出を試み、低加熱率においては温度測定値のばらつきが、高加熱率の場合は

熱物性値の不確かさが総合的な不確かさに対して支配的であることがわかった。本試験は、航技研極超音速

関連風洞対応風試の一環として実施されたものである。また本試験データは極超音速流におけるCFDコード

検証においても有用である。

properties and tunnel specifications. Specifically, HWT is a

blowdown type wind tunnel and therefore its flow properties can

be estimated relatively accurately. But the stagnation enthalpy

attainable in HWT is approximately 1 MJ/kg and is much lower

than actual flight conditions at hypersonic speeds. On the con-

trary, in HIEST, although much higher enthalpy levels up to 25

MJ/kg can be obtained, there are many unknown factors inherent

to the high-enthalpy short duration facilities. Hence it is mean-

ingful to use these facilities in a synergistic way, such that the

strength of one facility compensates for a weakness of the other.

In other words, one facility should be used to improve capability

and/or accuracy of the other such that both facilities benefit.

From this viewpoint, in JAXA, a comparative test program was

performed for the above two facilities to obtain force/heat flux

data using the same model configuration. Through the synergistic

use of these facilities together with the aid of CFD prediction

technique, we expect to have a practical guideline for accurate

and reliable prediction of aerothermodynamic properties of actu-

al flight vehicles.

In the present report, results are presented for the force/heat

flux measurement test conducted in the HWT using a ballistic-

type model configuration. The tests were conducted at conditions

of the stagnation enthalpy around 1 MJ/kg and the stagnation

pressures from 1 to 6 MPa. For the force test, the three-compo-

nent aerodynamic force data are compared with existing experi-

mental results obtained at other hypersonic facilities. To obtain

heat flux data, two measurement techniques were applied and the

results are compared to assess quantitative reliability of the data.

An uncertainty analysis was also performed to evaluate the mea-

surement uncertainty both for the force and heat test. In the force

test, a systematic approach to estimate both bias and random

errors for each experimental stage based on the statistical view-

point was used. In turn, since the data reduction includes a

numerical integration procedure for the heat test, a sophisticated

method to apply Monte Carlo technique was adopted to estimate

overall uncertainty of the heat flux reduction process. The pre-

dicted uncertainties are then compared with random errors esti-

mated from repeat tunnel runs of the present force/heat test.

In this report, the experimental data and corresponding free-

stream conditions are tabulated in detail as appendix and there-

fore this report is expected to be useful for the purpose of vali-

dating hypersonic CFD codes.

2. FACILITY

A schematic view of HWT is shown in Fig. 1. HWT is a blow-

down cold type wind tunnel with nominal Mach number of 10. To

prevent liquefaction of air, a pebble bed heater heated by a line

gas burner is utilized. The humidity management system includ-

ing the combustion gas replacement system keeps the humidity in

the working gas below 50 ppmV at a reservoir pressure of 4 MPa.

A previously conducted flow calibration test showed that the

Mach number uncertainty is less than 0.3% in the core flow part.

The basic test procedure of HWT is as follows. After the start

of blowdown, a model is injected into or withdrawn out of the

flow by a model support system in the force measurement and by

a rapid injection system in the heat flux measurement, respec-

tively. The main model support system is capable of changing the

model pitch angle continuously (sweep operation) or step-wise

(pitch-and-pause operation) during a tunnel run. The rapid injec-

tion system enables the model to be situated in the center of the

flow within 0.2 seconds from out of the flow so that the assump-

tion of step heating, required to reduce the heat flux from the IR

camera data using the one-dimensional heat conduction theory,

becomes appropriate. A detailed description of HWT facility is

found in Ref. [1].

3. MODELS

3.1 Model configuration

In the present study, a relatively simple model configuration

was employed because it is preferable to minimize uncertainties

coming from the geometry complexity from the viewpoint of tun-

nel-to-tunnel comparison. The HB-2 type model employed in the

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 3

Fig. 1 JAXA 1.27 m Hypersonic Wind Tunnel (HWT).

Fig. 2 HB-2 configurations.

4.9D

D

4.0D

10

0.3D

0.7D

25

1.6D

1.95D

Moment Reference

present test is a standard model proposed in a joint program of

AGARD and Supersonic Tunnel Association (STA) in

1950–60’s [2]. This has an analytical shape that consists of a

sphere, cone, cylinder, and flare as shown in Fig. 2.

3.2 Force model

The force model used in the present study is shown in Fig. 3 a)

and b). The model is made of a stainless steel (SUS304) and the

weight is 12.675 kg without the balance. The length and the cen-

terbody diameter of the model are 490 mm and 100 mm, respec-

tively. A six-component balance (Nissho LMC-6522-33/Z100)

was installed inside the model, and the balance center was set to

be located at the center of the model.

From a previously performed force test using the same model,

it has been appeared that the increase of balance temperature was

less than 1 K during the flow duration time and therefore the

effect of temperature drift on the balance is thought to be negligi-

ble. On the other hand, in order to perform base pressure correc-

tion, a total of three pressure sensors (Kulite Semiconductor

XCS-093-5A) were installed on the model base plate (two points)

and inside the model (one point). A detailed schematic of the

model is illustrated in Fig. A1 of Appendix A.

3.3 Heat model

The heat model are shown in Fig. 4 a) and b). The dimension is

the same as that of the force model. In this model, a total of 28

chromel-constantan type co-axial thermocouples of 1.5 mm

diameter (Medtherm TCS-E-10370) were press-fitted. The posi-

tion of each sensor is illustrated in Fig. A2 and Table A1 of

Appendix A.

A total of 8 sensors are circumferentially located in the flare

part to quantify uncertainties coming from the model align-

ment/flow deflection error at zero angle of attack. The model is

made of chromel in order to avoid electromotive force caused by

the difference of thermoelectric properties between the material

of the model surface and the outer tube of the thermocouple [3].

JAXA Research and Development Report JAXA-RR-04-035E4

Fig. 4 HB-2 heat model.(a) Chromel nose parts(b) Polyimide nose parts

(a)

(b)

Fig. 3 HB-2 force model.(a) Model close-up view(b) With model support system

(a)

(b)

The nose parts (29% of the total length, see Fig. 4) can also be

changed with Dupont Vespel polyimide plastic part which is

used for the non-intrusive surface temperature measurement

using infrared (IR) thermography technique. Vespel was selected

as a material since it is suitable from the viewpoint of homogene-

ity, low thermal conductivity, machinability, and decay durability

against the high temperature environment.

4. EXPERIMENTAL PROCEDURE

4.1 Test conditions

In Table 1, four standard operating conditions of HWT are tab-

ulated. The stagnation enthalpy is nearly constant irrespective of

the change in the stagnation pressure and is approximately 1

MJ/kg. The unit Reynolds number ranges from 0.9× 106 to

4.3× 106 /m.

As already mentioned, for the force test, two model injection

methods of the sweep and pith-and-pause operation can be select-

ed. The angle of attack ranges from －10 to 32 degrees for both

operations. For the sweep operation mode, the sweep angle was

divided into two parts (α = －10～ 18 and 16～ 32 degrees) due

to the limitation of flow duration time. For the pitch-and-pause

operation, data were obtained at five angles of attack (α = －10,

0, 10, 20, and 30 degrees) and the pause period was set to 2.5～

3.0 seconds at each angle of attack. A total of five repeat runs

were performed in the case of P0 = 1 MPa to estimate random

error from a statistical uncertainty analysis.

For the heat test, two angles of attack (0 and 15 degrees) were

selected. Similar to the force test, five repeat runs were performed

in the case of P0 = 2.5 MPa to estimate the random error. As well,

the model stream-wise location in the test section was changed to

evaluate uncertainties coming from stream-wise flow non-unifor-

mity. This was done by shifting the model injection position from

the standard position (500 mm from the nozzle exit) to 230 mm

upstream or to 450 mm downstream. For the nose part of the

model, a non-intrusive measurement using the infrared thermog-

raphy was also conducted.

Photographs of the rapid injection system and the infrared

camera placed in the test section are provided in Figs. 5 and 6,

respectively.

The list of run schedules and corresponding free-stream condi-

tions are summarized in Table C1 of the Appendix C for the force

test and C2 for the heat test, respectively. Care was taken not to

perform consequent runs on the same day. For example, each

repeat run was made one or two days apart to avoid comparing

two consecutive runs on the same day. This aids in estimating the

overall measurement system repeatability [4].

4.2 Data acquisition/reduction

As for the force test, the three-component balance outputs cor-

responding total axial force, normal force, and pitching moment

were converted to the aerodynamic coefficient by dividing the

dynamic pressure and reference area (plus reference length for

the pitching moment coefficient).

The loads acting on the balance are a combination of the aero-

dynamnic loads and the weight of the model. To extract the aero-

dynamic loads, the balance output must be corrected to remove

the effects of the model weight, which is termed as static weight

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 5

Fig. 6 Infrared camera system.

Fig. 5 Rapid injection system.

Table 1 HWT standard tunnel and free-stream conditions.

P0 (MPa) 1 2.5 4 6T0 (K)

M∞ 9.46870-970 920-1020 950-1030 1000-1070

9.59 9.65 9.69p∞ (Pa) 33 75 115 168T∞ (K) 51 52 53 55ρ ∞ (g/m3) 2.3 5.0 7.6 10.6

Re (× 106 1/m) 0.9 2.1 3.1 4.3

tare. The static tare data were obtained at the beginning of the test

campaign without blowing (wind-off), and were subtracted from

the wind-on data to correct the variation of the model weight load

with considering the sting/balance bending effects throughout the

run.

In a hypersonic wind tunnel testing, the static pressure is too

low to measure accurately and it is not practical to determine the

free-stream Mach number from the nozzle wall pressure. There-

fore the Mach number is computed from the stagnation pressure

and temperature measured at the reservoir together with the pitot

pressure measured in the test section. Since the pitot pressure

sensor cannot be placed simultaneously with a model, in HWT, a

calibration test using a number of pitot pressure sensors was con-

ducted and the Mach number was computed in advance. The

dynamic pressure was then calculated from the calibrated Mach

number and the stagnation pressure and temperature measured at

each run of the present force and heat tests. A detailed description

concerning the method of computing free-stream properties is

found in Appendix B.

The total axial force coefficient CA is converted to the forebody

axial force coefficient CAF through the base pressure correction as

CAF = CA－ (1)

where A and Ab are the reference area and the area of the model

base, respectively. On the other hand, the pitching moment coef-

ficient is first evaluated around the balance center, and is trans-

ferred to the value around the moment reference point as

Cm = CBm－ CN (2)

where d is the distance from the balance center to the moment

reference point (see Figs. 2 and A1).

For the heat flux measurement test, time history of the sensor

voltage was recorded in a data recorder (Yokogawa analysing

recorder AR 4800) for 5 seconds starting from the model injec-

tion time. A cold junction unit (Chino KT-C020) was utilized to

set the reference temperature. The voltage data were recorded at

a sampling frequency of 1 kHz and were smoothed by the moving

average technique. The smoothed sensor voltage was then con-

verted to temperature using a polynomial expression which

relates the thermoelectric voltage to the temperature. To reduce

aerodynamic heating from the thermocouple data, it is assumed

that 1) heat conduction along the body surface is negligible, and

2) thermal properties are not dependent on temperature. With

denoting the temperature measured at each time step (t0, t1, . . . tn)

as (T(t0), T(t1), . . . T (tn)), the model surface heat flux at time tn is

evaluated by applying the formula of Cook and Felderman [5],

[6], which is written as

(3)

where T’ denotes temperature increment from the initial value,

i.e., T’(ti) = T (ti) － T (t0). As for the thermal properties, a con-

stant chromel value at 300 K (ρck = 6.395 × 107 J2/m4K2s) was

taken from Ref. [7]. Finally, the computed time history of the heat

flux was averaged in time to reduce a time-averaged value for

each tunnel run.

For the IR thermography tests, the IR camera measurement

system was used. The IR camera measurement system consists of

an infrared camera (AGEMA 900LW) and a computer equipped

with a digitizer. The infrared image obtained was allocated to the

body surface through the three-dimensional image mapping. In

reducing aerodynamic heating from the image mapping data, the

method of Jones and Hunt was applied [8]. In this method, by

assuming a constant heat transfer coefficient h and step heating,

the analytical solution of the one-dimensional heat conduction

equation at time t (after the start of step heating) is available as

= 1 – exp(β2)erf(β ) (4)

where

(5)

The parameter β can be computed numerically from Eq. (4) by

specifying the initial temperature T(t0), measured temperature at

time t, T (t), and the adiabatic temperature Taw (assumed to be

equal to the tunnel stagnation temperature T0). The heat transfer

coefficient can then be obtained from Eq. (5). In this method,

only an initial temperature and a temperature at a time during a

tunnel run are required; i.e., no temperature time history is need-

ed. This is important for an imaging method since a large amount

of data space is required to store image files of the temperature

time history [9].

Also in this method, it is assumed that the thermal properties ρ,

c, and k are constant with respect to temperature. To correct the

temperature dependence, an effective temperature TJH was intro-

duced which is given as

TJH = T(t0)＋[T (t)－T(t0)]× F (6)

The thermal properties were then evaluated at this reference tem-

perature using a Vespel thermal properties curve fit. The factor F

β = ht

ρck

T(t) – T(t0)Taw – T(t0)

q(tn) = 2T’(ti) – T’(ti－ 1)tn－ ti＋ tn－ ti－ 1

n

Σi＝ 1

ρckπ

dD

(p∞ – pb) Ab

q∞A

JAXA Research and Development Report JAXA-RR-04-035E6

was empirically set to 0.6 and it has been shown that the effect of

thermal properties variation can be corrected within 1% accuracy

by using this value.

In a strict sense, as the model passes through the test section

wall boundary layer, the assumption of step heating does not hold

true. Therefore the model injection process was modeled as step

heating by correcting the time t in Eqs. (4) and (5) so that the heat

transfer coefficient linearly increases from zero to h and become

constant in the middle part of the wall boundary layer.

5. UNCERTAINTY ANALYSIS

5.1 Force test uncertainty

One conventional approach to estimate the accuracy of mea-

surement data is to assume the error (defined as the difference

between the experimentally determined value and truth) to be

composed of two components, namely bias and random errors.

The random error is defined as an uncertainty coming from the

scattering of the data, while the bias error is a systematic error

which is invariant throughout the test. Practically, both compo-

nents are to be quantified at each experimental process (i.e., cali-

bration, data acquisition, and data reduction). These components

are then summed up to evaluate the overall uncertainty at a spe-

cific level of confidence. The 95% confidence uncertainties for

the present force test were estimated using the methodology

described in Refs. [10] and [11].

To estimate the bias and random errors, a bias limit B and a

precision limit P are defined. Each limit is estimated by an inter-

val within which the true value of a variable lies. Specifically, the

precision limit is given as

P = tS (7)

where t is called the Student t value and can be determined

from the degree of freedom which is a measure of data indepen-

dency. The value S is the standard deviation of a sample of N

readings for a variable x, and S2 is called the unbiased estimate of

population variance which is defined as

S2 = (xk－–x )2 / (N－1) (8)

and the mean value –x is defined as

–x = xk / N (9)

Finally the 95% confidence uncertainty U is given by the root-

sum-square of the bias and precision limit as

(10)

In the present case, we need to estimate the uncertainty concern-

ing the aerodynamic coefficients CAF , CN, and Cm. Of the three,

the expression of the forebody axial force coefficient CAF found

in Eq. (1) is rewritten as

CAF = (11)

Hence the possible error includes uncertainties concerning the

axial aerodynamic force Fx, free-stream pressure p∞, base pres-

sure pb, and free-stream dynamic pressure q∞.

In the HWT experiment, the dynamic pressure is computed

from the following relation

q∞ = p∞M 2∞ (12)

where γ = 1.4 is the specific heat ratio. Note that the free-stream

Mach number is not dependent on each run but is a specified

value which was determined by the calibration test conducted in

advance. The free-stream static pressure p∞ is also reduced from

the calibrated Mach number with the tunnel stagnation pres-

sure/temperature measured at each tunnel run. Thus the uncer-

tainty in q∞ should be estimated as a combination of each ele-

mental error for M∞, P0, and T0.

As a first step, the uncertainty in the free-stream Mach number

M∞ is estimated. As described in Appendix B, the local Mach

number Mi obtained at each point in the test section is a function

of the tunnel stagnation pressure P0, stagnation temperature T0,

and pitot pressure Ppit. Table 2 summarizes the measurement

uncertainties of the tunnel properties.

Both bias and precision limits of the tunnel stagnation pressure

were reduced from calibration data of the pressure sensor, while

both error components were empirically determined for the stag-

nation temperature. The pitot pressure was measured at each

local point in the test section using electronically scanned pres-

sure (ESP) modules and hence its uncertainties were estimated

from calibration data of the ESP module.

Using these three measured properties, the Mach number at

each local point is computed based on the one-dimensional isen-

tropic relations with a caloric gas imperfections correction

method found in Ref. [12]. The concrete procedure of the Mach

number reduction method is described in Appendix B. Thus the

bias and precision limits concerning the Mach number uncertain-

ty can be estimated as a combination of each error component

BMi= [( BP0)

2

＋( BT0)2

＋( BPpit)2

]1/2

(13)∂Mi

∂Ppit

∂Mi

∂T0

∂Mi

∂P0

γ2

Fx－(p∞－ pb) Ab

q∞A

U = B2＋ P2

N

Σk＝1

N

Σk＝1

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 7

PMi= [( PP0)

2

＋( PT0)2

＋( PPpit)2

]1/2

(14)

The partial derivatives found in the above equations are called

sensitivity coefficients which represent the contribution of each

error component on the overall uncertainty. Since the Mach num-

ber reduction process includes a nonlinear operation, it is impos-

sible to analytically obtain these derivatives. Thus, in the present

case, they were evaluated by a numerical differentiation tech-

nique.

The free-stream Mach number M∞ is given as an average of the

local values as

M∞ = Mi /N (15)

where N is the number of measurement point in the test section.

Hence the free-stream Mach number uncertainty includes a com-

ponent concerning the scattering of each sensor data in addition

to the bias/random contributions of the ESP module. The result-

ing Mach number uncertainties are shown in Table 2.

Now the bias and precision limits for the dynamic pressure are

estimated as

Bq∞ = [( BP0)2

＋( BT0)2

＋( BMpit)2

]1/2

(16)

Pq∞ = [( PP0)2

＋( PT0)2

＋( PMpit)2

]1/2

(17)

In this case, the sensitivity coefficients are evaluated analytically

by differentiating Eq. (12) with respect to P0, T0, and M∞. How-

ever, since p∞ in Eq. (12) is also a function of P0, T0, and M∞, the

process is somewhat complex and hence they were also evaluat-

ed numerically in the present analysis. The estimated dynamic

pressure uncertainties are tabulated in Table 2.

In the next step, the uncertainty in the aerodynamic force mea-

surements are considered. The uncertainty in the balance output

consists of error components concerning the balance calibration

and the strain amplifier adjustment using a calibration strain gen-

erator. The level of these error sources were evaluated for both

bias and precision limits from cataloged data. The total uncer-

tainty was found to be around 0.1% of the balance capacity for

each force component. As mentioned in the preceding section, in

the reduction of aerodynamic forces, the static tare data obtained

before the experiment are subtracted from the balance output for

each angle of attack. Thus the aerodynamic force at each angle of

attack includes uncertainties concerning the static tare measure-

ment, wind-on force measurement, and angle of attack interpola-

tion. Both for wind-on and static tare measurements, the bias

limit is equal to that of the balance output, while the scattering of

the data during the measurement was added to evaluate the preci-

sion limits. The aerodynamic force uncertainties coming from the

angle of attack interpolation were reduced from the estimated

angle of attack error. The resulting bias and precision limits of

these error sources are combined to calculate the total uncertain-

ty in the aerodynamic force measurements in Table 3.

Finally the aerodynamic forces Fx and Fz are converted to the

aerodynamic coefficients CAF and CN by dividing the dynamic

pressure and the reference area. In the present case, the base pres-

sure contribution to the overall uncertainty in CAF was found to be

negligible and omitted. In this case, the overall uncertainty in CAF

can be estimated as

BCAF= [( BFx)

2

＋( Bq∞)2

]1/2

(18)∂CAF

∂q∞

∂CAF

∂Fx

∂q∞∂M0

∂q∞∂T0

∂q∞∂P0

∂q∞∂M0

∂q∞∂T0

∂q∞∂P0

N

Σi＝1

∂Mi

∂Ppit

∂Mi

∂T0

∂Mi

∂P0

JAXA Research and Development Report JAXA-RR-04-035E8

Typical value 1000000 2500000 4000000 6000000B 207 207 207 587S 1533 1533 1533 6212U 3422 3422 3422 14066

Typical value 920 970 990 1030B 30 30 30 30S 5 5 5 5U 31.7 31.7 31.7 31.7

3578 8945 14312 21468B 3.1 3.1 3.1 3.1S 5.5 5.5 5.5 5.5U 11.6 11.6 11.6 11.6

9.46 9.59 9.65 9.69B 0.011 0.010 0.010 0.010S 0.014 0.014 0.013 0.011U 0.029 0.030 0.028 0.025

2135 5010 7788 11459B 16 37 55 79S 15 35 50 65U 34 78 113 149

Stagnation pressure (Pa)

Stagnation temperature (K)

Pitot pressure (Pa)

Mach number

Dynamic pressure (Pa)

Typical value

Typical value

Typical value

Table 2 Estimated uncertainties in the HWT tunnel and free-stream properties.

PCAF= [( PFx)

2

＋( Bq∞)2

]1/2

(19)

UCAF= [(BCAF

)2＋(PCAF)2]1/2 (20)

This is possible since there is no correlation between Fx and q∞.

If the base pressure correction term is included, the partial deriv-

ative of CAF should be evaluated with respect to the independent

variables P0, T0, and M∞ instead of q∞. The partial derivatives

∂CAF /∂q∞ and ∂CAF /∂Fx can be evaluated readily by directly dif-

ferentiating Eq. (11) with respect to Fx and q∞. The uncertainty

in the normal force coefficient CN can be reduced by the same

procedure.

On the other hand, the pitching moment coefficient around the

moment reference point Cm is computed from the pitching

moment around the balance center CBm as

Cm = CBm－ CN (21)

Hence the bias and precision limits for the pitching moment coef-

ficient are evaluated as

BCm= [( BCB

m)2

＋( Bd)2

＋( BcN)2

＋ 2 BCBm

BCN]1/2

(22)

PCm= [( PCB

m)2

＋( Pd)2

＋( PCN)2

]1/2

(23)

UCm= [B2

Cm＋ P2

Cm]1/2 (24)

The bias limit includes a cross term which correlates CBm and CN

since the uncertainties of the pitching moment and the normal

force arise from the same source and are presumed to be perfect-

ly correlated.

The resulting predicted uncertainties for each aerodynamic

force coefficient are tabulated for the case of α = 15 degrees in

Table 4.

5.2 Heat test uncertainty

As described previously, the heat flux is evaluated from the

temperature time history of the thermocouples by using Eq. (3),

which is rewritten as

(25)

Therefore the accuracy of computed heat flux is influenced by

uncertainties in the temperature measurement and in the thermal

properties of a material. It should be noted that only the tempera-

ture increment, T’(ti) = T(ti)－ T (t0), is required to evaluate the

heat flux from Eq. (25) and no magnitude of measured value is

necessary. In this case, a series of bias errors are expected to be

offset and only the random component is required to be account-

ed for. To properly estimate the precision limit of the heat flux

data, in the present study, a Monte Carlo algorithm was newly

developed. In this algorithm, the temperature data were random-

ly varied by a specified level of uncertainty and the overall uncer-

q(tn) = 2T’(ti) – T’(ti－ 1)tn－ ti＋ tn－ ti－ 1

n

Σi＝ 1

ρckπ

∂Cm

∂CN

∂Cm

∂d

∂Cm

∂CBm

∂Cm

∂CN

∂Cm

∂CBm

∂Cm

∂CN

∂Cm

∂d

∂Cm

∂CBm

dD

∂CAF

∂q∞

∂CAF

∂Fx

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 9

Table 3 Estimated uncertainties in the aerodynamic force measurement.

Table 4 Estimated uncertainties in the aerodynamic coefficientsat angle of attack of 15 degrees

tnenopmoC Fx )N( Fy )N( Fz )N( M x )mN( M y )mN( M z )mN(elacslluF 542 094 189 51 47 94

tnemerusaemeratcitatSB 340.0 790.0 02.0 700.0 420.0 510.0S 52.0 04.0 53.0 310.0 820.0 020.0U 94.0 97.0 17.0 620.0 060.0 140.0

tnemerusaemno-dniWB 340.0 790.0 02.0 700.0 420.0 510.0S 42.0 04.0 43.0 310.0 920.0 020.0U 84.0 08.0 96.0 620.0 160.0 140.0

kcattafoelgnAB 840.0 120.0 310.0 000.0 000.0 000.0S 740.0 020.0 310.0 000.0 000.0 000.0U 11.0 640.0 920.0 000.0 000.0 100.0

latoTB 80.0 41.0 92.0 010.0 330.0 120.0S 53.0 65.0 84.0 810.0 040.0 820.0U 96.0 21.1 99.0 730.0 680.0 850.0

P0 1 2.5 4 6

CAF

0.767 0.723 0.722 0.712B 0.007 0.004 0.004 0.005S 0.021 0.009 0.006 0.006U 0.043 0.018 0.013 0.012

CN

1.175 1.140 1.174 1.169B 0.019 0.009 0.008 0.009S 0.030 0.013 0.009 0.008U 0.062 0.027 0.019 0.019

Cm

-1.092 -1.050 -1.098 -1.090B 0.030 0.013 0.010 0.009S 0.028 0.013 0.009 0.007U 0.063 0.028 0.020 0.016

(MPa)Typical value

Typical value

Typical value

tainty in the time-averaged heat flux was numerically estimated.

By taking into account the data reduction process employed in

the actual test, the random uncertainty component coming from

the temperature measurement error was evaluated based on the

following procedures.

1. Set a constant value of the surface heat flux q0.

2. Obtain a baseline temperature time history T0,i by the fol-

lowing relation given from the exact solution of the one-

dimensional heat conduction equation [6]

ti = i∆t

where ∆t is the time increment which corresponds to the

data sampling frequency.

3. Set l = 0, where l is the number of iteration for the step

4 to 8.

4. Set the “noisy” temperature time history Ti as

Ti = T0,i＋ ∆T(2U－1)

where U is the random variable which takes a value

between 0 and 1, and ∆T is the maximum level of random

uncertainty.

5. Smooth the temperature time history by the moving aver-

age.

6. Obtain the heat flux time history from Eq. (25) for a set of

“smoothed” Ti.

7. Compute time-averaged heat flux qlav.

8. Evaluate the variance as

σ 2 = (qiav－q0)

2/l

9. Increment l as l→ l＋ 1 and repeat the step 4 to 8 until the

level of the heat flux uncertainty (2σ for 95% coverage)

converges to a constant value.

Care was taken such that each of the above uncertainty estimation

process consistently follows the actual data reduction process. In

the present study, the temperature random uncertainty ∆T was

estimated from the experimental data as a standard deviation of

the temperature scattering and was set to 0.2 degrees. The heat

flux uncertainty corresponding to the specified ∆T was then com-

puted by varying the level of surface heat flux from 1 to

100 kW/m2. Note that the heat flux level of 1 and 100 kW/m2 is

the same order as the present experimental data obtained at the

cylinder/flare and stagnation part of the model, respectively.

On the other hand, since the term concerning the thermal prop-

erties appears as a constant in Eq. (25), the thermal properties

uncertainty was estimated as a bias limit. The coefficient is

rewritten as , where α denotes thermal diffusivity. From

the previous experience, the measurement uncertainty of ρ, c, and

α were considered to be 1%, 1%, and 5%, respectively. Hence the

uncertainty of can be evaluated as 2.9%.

In summary, the overall heat flux uncertainty was estimated as

a root-sum-square combination of the random component con-

cerning the temperature measurement and the bias component

concerning the thermal properties estimation. In Fig. 7, the esti-

mated heat flux uncertainty as a function of the heat flux level is

depicted.

It can be confirmed that, in the case of q0 = 1 kW/m2, the total

uncertainty reaches up to 40%. As expected, the effect of temper-

ature uncertainty becomes dominant as the level of heat flux

decreases, while becomes negligible for the case of q0 =

100 kW/m2. The data smoothing reduces the heat flux uncertain-

ty to a large extent for the low heating case. In fact, as much as

97% error was observed for the q0 = 1 kW/m2 case if the data

were not smoothed. In this case, the temperature increase is of the

order of 0.1 K, which is even smaller than the level of the tem-

perature scattering specified. Hence, although the data are suffi-

ciently smoothed by moving average, we cannot completely get

rid of the random errors. On the contrary, for the high heating

case, up to several ten degrees of temperature increase is noted.

Hence the effect of temperature random uncertainty becomes

negligible and we can obtain essentially identical results even

though the data are not smoothed. Therefore the total heat flux

estimation accuracy is essentially affected by only the level of

uncertainty in the thermal properties.

ρck

ρc α

ρck

l

Σi＝1

T00

,i = , i = 0, . . . , nti

ρck2qπ

JAXA Research and Development Report JAXA-RR-04-035E10

Heat Flux (kW/m2)

)%( ytniatrecn

U

TemperatureThermal PropertiesTotal

1 10 1000

10

20

30

40

Fig. 7 Estimated heat flux errors versus the heat flux level.

6. RESULTS AND DISCUSSION

6.1 Force tests

The variation of the three-component aerodynamic coeffi-

cients CAF, CN, and Cm with respect to the angle of attack are

shown in Figs. 8 to 15 for the nominal stagnation pressures of 1,

2.5, 4, and 6 MPa, respectively.

For the forebody axial force coefficient, the contribution of the

base pressure correction term (see Eq. (1)) to the total value is

shown to have, for example, a maximum of 4% for the case of P0

= 6 MPa and hence cannot be neglected.

In Figs. 8, 10, 12, and 14, the predicted levels of uncertainty in

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 11

Fig. 8 Forebody axial force coefficient versus angle of attack(P0 = 1 MPa).

Fig. 10 Forebody axial force coefficient versus angle of attack(P0 = 2.5 MPa).

Fig. 11 Normal force and pitching moment coefficients versusangle of attack (P0 = 2.5 MPa).

Fig. 12 Forebody axial force coefficient versus angle of attack(P0 = 4 MPa).

Fig. 13 Normal force and pitching moment coefficients versusangle of attack (P0 = 4 MPa).

Fig. 9 Normal force and pitching moment coefficients versusangle of attack (P0 = 1 MPa).

Angle of Attack (deg)

C ,tneiciffeoC ecroF laix

A ydoberoFF

A Run 1465, 1470, 1473, 1481, 1483Run 1468, 1471Run 1469, 1472 (pitch and pause)95% Uncertainty (predicted)

-10 0 10 20 30

0.6

0.8

1

1.2

Angle of Attack (deg)

C ,tneiciffeoC ecroF la

mroN

N

C ,tneiciffeoC tne

moM gnihcti

Pm

CN

Cm

Run 1465, 1470, 1473, 1481, 1483Run 1468, 1471Run 1469, 1472 (pitch and pause)

-10 0 10 20 30

0

1

2

3

4

-4

-3

-2

-1

0

Angle of Attack (deg)

C ,tneiciffeoC ecroF laix

A ydoberoFF

A Run 1462, 1467Run 1463, 148095% Uncertainty (predicted)

-10 0 10 20 30

0.6

0.7

0.8

0.9

1

1.1

Angle of Attack (deg)

C ,tneiciffeoC ecroF la

mroN

N

C ,tneiciffeoC tne

moM gnihc ti

Pm

CN

Cm

Run 1462, 1467Run 1463, 1480

-10 0 10 20 30

0

1

2

3

4

-4

-3

-2

-1

0

Angle of Attack (deg)

C ,tneiciffeoC ecroF laix

A ydoberoFF

A Run 1460, 1466Run 1461, 148295% Uncertainty (predicted)

-10 0 10 20 30

0.6

0.7

0.8

0.9

1

1.1

Angle of Attack (deg)

C ,tneiciffeoC ecroF la

mroN

N

C ,tneiciffeoC tne

moM gnihcti

Pm

CN

Cm

Run 1460, 1466Run 1461, 1482

-10 0 10 20 30

0

1

2

3

4

-4

-3

-2

-1

0

CAF are also indicated. Although the level of uncertainty varies as

the angle of attack changes, the minimum value at zero degrees

was used for all angles of attack. The level of uncertainty for the

case of P0 = 1 MPa reaches up to 6% of the measurement value

and decreases as the stagnation pressure increases. As far as the

forebody axial force coefficient is concerned, the error element

concerning the balance measurement is dominant for the overall

uncertainty and hence the level of uncertainty increases as the

aerodynamic force decreases, i.e., as the stagnation pressure

decreases. For each level of the stagnation pressure, it can be con-

firmed that all data agreed well within the predicted uncertainty.

Although a slight discrepancy is observed at the overlapped

angles of attack between the low- and high-sweep operation

mode, its discrepancy is much smaller than the level of uncer-

tainty. As well, two pitch-and-pause data differs slightly as indi-

cated in Fig. 8, but the difference is within the level of uncertain-

ty range and is hence considered to be acceptable.

Table 5 shows a comparison of the estimated/experimentally-

evaluated precision limits of the three-component aerodynamic

coefficients in the case of P0 = 1 MPa and an angle of attack of 15

degrees. The experimental values were obtained from a total of

five repeat runs with applying Eqs. (7) to (9).

Overall, the estimated precision limit is larger than the experi-

mental value. This is presumably due to overestimation of the

balance output errors.

For the case of P0 = 6 MPa, the results are compared with data

obtained at Arnold Engineering Development Center (AEDC)

50-inch Mach 10 Tunnel [2] as shown in Fig. 15. In this tunnel,

the Reynolds number based on the centerbody diameter of the

model is 1.36× 106, which is much higher than the present HWT

conditions (see Table C1). Excellent agreement is confirmed for

the normal and pitching moment coefficients.

In Fig. 16, the zero-lift forebody axial force coefficient is plot-

ted versus the viscous parameter defined as M/ . At hyperson-

ic speeds, the wave drag (pressure integration) contribution to the

total aerodynamic coefficient remains constant with Reynolds

number variations, while the low local Reynolds number, which

is caused by the increased bow-wave total pressure losses, pro-

duces a relatively large skin friction contribution [2]. Hence the

axial force increases as the Reynolds number decreases.

Finally the three-component aerodynamic coefficients versus

the angle of attack for each tunnel run are tabulated in detail in

Table D1 of Appendix D.

Re

JAXA Research and Development Report JAXA-RR-04-035E12

Fig. 14 Forebody axial force coefficient versus angle of attack(P0 = 6 MPa).

Fig. 15 Normal force and pitching moment coefficients versusangle of attack (P0 = 6 MPa).

Fig. 16 Zero-lift axial-force coefficient versus viscous parame-ter.

Table 5 Comparison of the predicted/experimentally-evaluatedprecision limits (P0 = 1 MPa, α = 15 deg).

Angle of Attack (deg)

C ,tneiciffeoC ecroF laix

A ydoberoFF

A Run 1477, 1486, 1488Run 1475, 1478, 148795% Uncertainty (predicted)

-10 0 10 20 30

0.6

0.7

0.8

0.9

1

Angle of Attack (deg)

C ,tneiciffeoC ecroF la

mroN

N

C ,tneiciffeoC tne

moM gnihc ti

Pm

CN

Cm

Run 1477, 1486, 1488Run 1475, 1478, 1487AEDC data

-10 0 10 20 30

0

1

2

3

4

-4

-3

-2

-1

0

Viscous Parameter, M/Re1/2C ,tneiciffeo

C ecroF laixA ydoberoF tfi

L-oreZ

FA

P0=1MPa P0=2.5MPa P0=4MPa P0=6MPa

0.015 0.02 0.025 0.03 0.035

0.56

0.6

0.64

0.68

CAF 0.021 0.011CN 0.030 0.018Cm 0.028 0.024

Estimated fromuncertainty analysis

Obtained from5 repeat runs

6.2 Heat tests

In Fig. 17, the heat transfer coefficient distributions along the

body surface obtained from the thermocouples are compared

between the five repeat runs for the case of P0 = 2.5 MPa at zero

angle of attack. It is noted that the heat transfer coefficient data

agree well in the nose part, while they are slightly scattered in the

cylinder/flare junction part. In this region, the temperature

increase is small due to relatively low heating and thus the S/N

ratio of sensor output becomes degraded. This is, however,

expected since up to 40% of uncertainty was evaluated at the heat

flux level of 1 kW/m2 in the preceding uncertainty analysis. On

the other hand, the random errors in the nose part is less than 1%

for every sensors and thus good repeatability is confirmed.

In Fig. 18, non-dimensional heat flux distributions (divided by

the stagnation heat flux value) along the body surface are com-

pared for the four levels of the stagnation pressure. As can be

seen, the trend of the distribution is nearly identical for each P0

and hence the effect of the reservoir pressure variation (i.e.,

Reynolds number) is found to be negligible in the present test

condition range.

The axial distributions of the heat transfer coefficient for the

model windward part at an angle of attack of 15 degrees are plot-

ted for the case of P0 = 2.5 MPa in Fig. 19 and for the comparison

between each P0 in Fig. 20, respectively. Compared to the zero

angle of attack case, the level of aerodynamic heating is relative-

ly high even in the cylinder/flare junction region and hence good

repeatability is confirmed throughout the whole part.

Figure 21 indicates the comparison of heat flux distributions in

the nose part between the data reduced from the co-axial thermo-

couples and those from the IR thermography. Good agreement is

confirmed between the two measurement techniques, typically

less than 3% discrepancy in the stagnation region. Since planar

measurement is possible for the IR thermography technique, the

data which have the same axial position were compared at zero

angle of attack conditions. The result showed that the data scat-

tering in the circumferential direction was typically less than 1%

and hence was supposed to be negligible. As well, a set of data

reduced from the different data acquisition time were compared

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 13

Fig. 17 Repeatability of the heat transfer coefficients (P0 =2.5 MPa, α = 0 degrees).

m/W( tneiciffeo

C refsnarT tae

H2

)K

x (mm)

Run 1323 (P0=2.5MPa)Run 1334 (P0=2.5MPa)Run 1339 (P0=2.5MPa)Run 1340 (P0=2.5MPa)Run 1343 (P0=2.5MPa)

0 100 200 300 400

1

10

100

Fig. 18 Comparison of non-dimensional heat flux distributionsfor different stagnation pressures (α = 0 degrees).

q/qgts

x (mm)

Run 1346 (P0=1MPa)Run 1323 (P0=2.5MPa)Run 1324 (P0=4MPa)Run 1322 (P0=6MPa)

0 100 200 300 400

0.01

0.1

1

Fig. 19 Repeatability of the heat transfer coefficients (P0 =2.5 MPa, α = 15 degrees).

Fig. 20 Comparison of non-dimensional heat flux distributionsfor different stagnation pressures (α = 15 degrees).

m/W( tneiciffeo

C refsnarT tae

H2

)K

x (mm)

Run 1335 (P0=2.5MPa)Run 1342 (P0=2.5MPa)Run 1347 (P0=2.5MPa)

0 100 200 300 400

10

100

q/qgts

x (mm)

Run 1335 (P0=2.5MPa)Run 1336 (P0=4MPa)Run 1337 (P0=6MPa)

0 100 200 300 400

0.1

1

and it appeared that the difference was also less than 1%. The

reliability of the present data is thus enhanced by applying the

two measurement techniques.

As already mentioned, a total of five repeat runs were made to

evaluate uncertainties associated with flow repeatability. The pre-

cision limit for the heat transfer coefficient was derived by using

Eqs. (7) to (9) and is illustrated for each sensor in Fig. 22 (data of

sensor 10 are not included due to sensor trouble). As expected, a

maximum uncertainty of up to 20% is noted around sensor 15,

corresponding to the cylinder/flare junction point, due to low

heating. For the case of P0 = 2.5 MPa, the minimum heat flux

level is around 1.5 kW/m2. It should be noted from Fig. 7 that the

predicted level of random uncertainty at this heat flux level is

almost the same, indicating that the present Monte Carlo analysis

estimates the heat flux uncertainty reasonably.

Next the effect of the test section flow non-uniformity on the

measurement accuracy was examined by changing the model

injection point and the result is shown in Fig. 23. The difference

is less than 1% in the nose part and is the same order of the data

repeatability as can be seen in Fig. 22. Therefore it appears that

there is no significant effect concerning the change of the model

injection location and, in other words, the flow field is sufficient-

ly uniform in the stream-wise direction.

Next the data are compared at the zero angle of attack condi-

tions by rotating the model around the body axis prior to the tun-

nel run. The rolling angle φ is defined as zero when the sensors in

the cylindrical part (e.g. sensor 15) face upward. Then the model

was rotated in the clockwise direction by 90, 180, and 270

degrees, viewed from the downstream. Two differences of the

heat transfer coefficient between φ = 180 and 0 degrees and

between φ = 270 and 90 degrees are shown for each sensor in

Fig. 24.

As confirmed, the sign of the error changes from minus to plus

at the sensor 4 (corresponds to the stagnation point). This implies

that the model is slightly inclined relative to the free-stream such

that sensors 1 to 3 (see Fig. A2 of Appendix A) faces upwind at φ

= 0° and 90° even though the nominal angle of attack is zero. This

is also confirmed from Fig. 25 which shows the distribution of

the averaged heat transfer coefficient obtained from the five

repeat runs for 8 sensors circumferentially placed in the flare

JAXA Research and Development Report JAXA-RR-04-035E14

Fig. 21 Comparison of heat transfer rate distributions betweenthermocouple and IR thermography data (P0 = 2.5 MPa,α = 0 degrees).

m/Wk( h

2)

K

x/D

Co-axial thermocoupleInfrarated thermography

Body shape

0 0.2 0.4 0.6 0.8 1 1.2 1.4

0.05

0.1

0.15

0.2

Fig. 22 Random error (precision limit) distribution obtainedfrom the five repeat runs (P0 = 2.5 MPa).

Fig. 23 Error distribution due to the change of model injectionpoint (P0 = 2.5 MPa).

Fig. 24 Error distribution due to the change of model rotationalangle (P0 = 2.5 MPa).

1 2 3 4 5 6 7 8 9 11 12 13 14 15 16 17 18 19 20 210

2

4

6

8

10

12

14

16

18

20

Precision limitP0 = 2.5MPa, T0 =700°C

Sensor No.

Err

or (

%)

1 2 3 4 5 6 7 8 9 11 12 13 14 15 16 17 18 19 20 210

2

4

6

8

10

12

14

16

18

20

Sensor No.

Err

or (

%)

|L2-L1||L3-L1|

Effect of mode linjection locationP0 = 2.5MPa, T0 =700°C

1 2 3 4 5 6 7 8 9 11 12 13 14 15 16 17 18 19 20 21

180° - 0°270° - 90°

0

4

8

12

16

20

24

Sensor No.

Effect of model rolling angleP0 = 2.5MPa, T0 =700°C

Err

or (

%)

part.

The error bars shown in Fig. 25 denote the precision index con-

cerning the averaged value. Considering the level of uncertainty,

it is obvious that there is a systematic difference such that the heat

transfer coefficient of the sensors 25 and 27 is much higher than

that of the sensors 21 and 23, respectively. Hence it can be con-

sidered that the model is inclined relative to the free-stream so

that sensors 25 and 27 face upwind. Considering the correspon-

dence of the sensor locations between the nose and flare part (see

Fig. A2), this is consistent with the result of Fig. 24. In fact, the

actual model alignment angle at the zero angle of attack condi-

tion was measured in the present experiment using a level and

0.16 degrees of the model support error was confirmed. Since the

degree of the scattering is found to be around 10%, we can con-

clude that the heat transfer coefficients in the flare part include

10% uncertainty produced by the model alignment error in addi-

tion to the random/bias errors contained in each sensor.

Finally, the present heat test data are summarized in detail for

each sensor and tunnel run in Table D2.

6. CONCLUSIONS

As a series of the comparison test campaign between two

hypersonic facilities in JAXA, force and heat flux measurements

were conducted in the JAXA 1.27 m blow-down cold type hyper-

sonic wind tunnel using the ballistic type standard model HB-2.

From the force test data obtained, good repeatability was con-

firmed for all of the tunnel conditions. The magnitude of mea-

surement uncertainty was estimated using a statistical approach

and it was observed that the scattering of the experimental data

were reasonably included within the predicted uncertainty limits.

The normal force and pitching moment data were also compared

with existing experimental data conducted in the AEDC hyper-

sonic wind tunnel, showing a good agreement.

As for the heat test, good repeatability in terms of the heat

transfer coefficient was also confirmed concerning the heat trans-

fer coefficient distribution along the model surface for the nose

part, while a slight scattering was observed in the vicinity of the

cylinder/flare junction part due to low heating. The precision

limit was evaluated from the data of five repeat runs and it was

found that the estimated value for the heat transfer coefficient is

less than 1% for the high-heating and up to 20% in the low-heat-

ing part, respectively. The heat transfer coefficient obtained from

the IR thermography agreed very well with the thermocouple

data. From the data for eight sensors installed circumferentially

in the flare part, it was shown that up to 10% of uncertainty in the

heat transfer coefficient could exist due to model support error.

Also performed was a Monte Carlo simulation to estimate the

level of uncertainty in the time-averaged heat flux data reduction

process. The result showed that the uncertainty in thermal prop-

erties is dominant for overall accuracy of the measurement data

when the heating rate is sufficiently high (e.g. 100 kW/m2). On

the other hand, for a heat flux level of 1 kW/m2, the effect of data

scattering concerning the temperature measurement was found to

become dominant. The predicted level of the precision limit in

the low heating part showed a good agreement with that experi-

mentally obtained from the five repeat runs.

Finally, since the data were carefully examined in terms of

accuracy and were confirmed to be sufficiently reliable, the pre-

sent force/heat test results are believed to provide useful informa-

tion as a database for the validation of hypersonic CFD codes.

ACKNOWLEDGMENTS

The authors would like to thank Mr. Shigeo Kayaba and Mr.

Muneyoshi Nakagawa for their contributions to the present test

campaign. One of the authors would also like to thank Dr.

Keisuke Fujii for the valuable discussion concerning the heat flux

reduction process using the co-axial themocouple.

REFERENCES

1) Nomura, S., Sakakibara, S., Hozumi K., and Soga, K.,

“NAL New Hypersonic Wind Tunnel System,” AIAA Paper

AIAA 93-5006, Dec. 1993.

2) Gray, J. D., “Summary Report on Aerodynamic Characteris-

tics of Standard Models HB-1 and HB-2,” AEDC-TDR-64-

137, July 1964. 25.

3) Kidd, C.T., Nelson C.G., and Scott, W.T., “Extraneous Ther-

moelectric EMF Effects Resulting from the Press-Fit Instal-

lation of Coaxial Thermocouples in Metal Models,” Pro-

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 15

Fig. 25 Comparison of heat transfer coefficient in the azimuthaldistribution at the flare part (P0 = 2.5 MPa).

4.4

4.6

4.8

5

5.2

5.4

5.6

m/W( tneiciffeo

C refsnarT tae

H2

)K

Sensor No.

2122

23

24

25

26

27

28

ceedings of the 40th International Instrumentation Sympo-

sium, Baltimore, MD, May 1994, pp. 317–335.

4) Aeschliman, D.P. and Oberkampf, W.L., “Experimental

Methodology for Computational Fluid Dynamics Code Val-

idation,” AIAA J., Vol. 36, No. 6, May 1998, pp. 733-741.

5) Cook, W. J., and Felderman, E. J., “Reduction of Data from

Thin-Film Heat-Transfer Gages: A Concise Numerical

Technique,” AIAA J., Vol. 4, No. 3, March 1966, pp.

561–562.

6) Schultz, D.L., and Jones, T. V., “Heat Transfer Measure-

ments in Short Duration Hypersonic Facilities,” AGARDo-

graph 165, 1973.

7) Sundqvist, B., “Thermal Diffusivity and Thermal Conduc-

tivity of Chromel, Alumel, and Constantan in the range

100.450K,” J. Appl. Phys., Vol. 72, No. 2, July 1992, pp.

539.

8) Jones, R. A., and Hunt, J. L., “Use of Fusible Temperature

Indicator for Obtaining Quantitative Aerodynamic Heat-

Transfer Data,” NASA TR R-230, 1966.

9) Merski, N. R., “Global Aeroheating Wind-Tunnel Measure-

ments Using Improved Two-Color Phosphor Thermography

Method,” AIAA J., Vol. 36, No. 2, March-April 1999, pp.

160–170.

10) Assessment of Wind Tunnel Data Uncertainty, AIAA Stan-

dard S-071-1995, AIAA, Washington, DC, 1995.

11) Measurement Uncertainty, ASME Performance Test Codes,

Supplement on Instruments and Apparatus, Part 1,

ANSI/ASME PTC19.1-1985, 1985.

12) Boudreau, A. H., “Performance and Operational Character-

istics of AEDC/VKF Tunnels A, B, and C,” AEDC-TR-80-

48, 1981.

13) Ames research staff, “Equations, Tables, and Charts for

Compressible Flow,” NACA Report 1135, 1954

Appendix A

Detail of the force/heat model

JAXA Research and Development Report JAXA-RR-04-035E16

490.0

100.0

Dimension in mm

Heat flux sensor

12

345

67 89

10 11 12 13 14 15 16 17 18 19 2021

28

26

25

160.027

21

22

23

24

25

26

27

28

View from backward

y

x

Fig. A1 Detailed drawing of HB-2 force model.

Fig. A2 Sensor location of HB-2 heat model.

Table A1 Heat flux sensor location.

Sensor No. x (mm) y (mm) θ (deg)1 13.9 -25.3 0.02 6.5 -18.6 0.03 1.7 -9.9 0.04 0.0 0.0 0.05 1.7 9.9 0.06 6.5 18.6 0.07 13.9 25.3 0.08 22.9 29.8 0.09 41.0 38.2 0.0

10 59.3 46.3 0.011 78.9 49.9 0.012 108.9 50.0 0.013 170.7 50.0 0.014 228.1 50.0 0.015 254.8 50.0 0.016 275.6 50.0 0.017 300.4 50.3 0.018 329.8 52.5 0.019 364.6 57.9 0.020 404.0 64.9 0.021 439.3 71.1 0.022 439.3 50.3 4523 439.3 0.0 9024 439.3 -50.3 13525 439.3 -71.1 18026 439.3 -50.3 22527 439.3 0.0 27028 439.3 50.3 315

Appendix B

Computation of free-stream properties in HWT

In reducing the free-stream properties for a conventional cold

type hypersonic wind tunnel, one approach accepted widely is to

assume an isentropic expansion from the reservoir to the test sec-

tion. In this case, the free-stream conditions can be computed

from the stagnation pressure P0, stagnation temperature T0, and

pitot pressure Ppit.

The following one-dimensional isentropic formulas provide

relations between the stagnation and free-stream flow properties

as functions of the Mach number [13]

( ) perfect=[ ] [ ] (B1)

( )perfect=[1＋ M2] (B2)

( )perfect=[1＋ M2]－1

(B3)

where γ = 1.4 is the specific heat ratio and the subscript “per-

fect” denotes perfect gas (i.e., “ideal” property). Note that these

relations hold only for perfect gas flows. In a hypersonic wind

tunnel, the stagnation pressure and temperature are high so that

effects of intermolecular force and vibrational energy excitation

become not negligible. To take into account such real gas effects,

the following formulas found in Ref. [12] give useful correlation

factors applicable for a range of calorically and thermally imper-

fect gases:

( ) real / ( ) perfect= 1.0562＋ 49.57 × 10－6P0

－(3523＋1.8300P0＋1.3839T0－0.0002196P0T0)×10－8T0

(B4)

( ) real / ( ) perfect= 0.9378－3.900 × 10－6P0

+ (6533＋0.6547P0－0.4137T0－0.0001354P0T0)×10－8T0

(B5)

( ) real / ( ) perfect= 1.0419＋ 38.31 × 10－6P0

－(1968＋ 0.7925P0＋ 1.6905T0)×10－8T0 (B6)

where the subscript “real” denotes real gas (i.e., “measured”

property) and the unit of P0 and T0 are psi and °R, respectively

(1 psi = 1/6894.76 Pa and 1 °R = 9/5 K). By converting the “real”

(measured) properties to the “perfect” (ideal) values, it becomes

possible to apply the perfect gas isentropic relations Eqs. (B1) to

(B3). The above correlation factors were calculated based on the

Beattie-Bridgeman equation of state with considering the vibra-

tional relaxation.

The local Mach number Mi found in Eq. (15) is computed by

the following procedure:

1. Evaluate (Ppit/P0)real from the measured P0 and Ppit.

2. Obtain (Ppit/P0)perfect from Eq. (B6).

3. Compute Mi from Eq. (B1). Since Eq. (B1) is a nonlinear

equation with respect to M, it should be solved using an

iterative technique.

Thus the free-stream Mach number M∞ is calculated as an aver-

age of the local Mach number obtained at each point in the test

section as shown in Eq. (15).

On the other hand, the free-stream conditions for the present

force/heat test were computed from P0 and T0 measured at each

tunnel run plus the Mach number obtained as above. The detailed

procedures are listed below.

1. Obtain (p∞/P0)perfect and (T∞/T0)perfect from the specified M∞

and Eqs. (B2) and (B3).

2. Obtain (p∞/P0)real and (T∞/T0)real from Eqs. (B4) and (B5).

3. Obtain p∞ and T ∞ as p∞ = (p∞/P0)real× P0 and T∞ =

(T∞/T0)real× T0.

4. Compute the dynamic pressure q∞ from the following rela-

tion.

q∞ = p∞M 2∞ (B7)

The free-stream conditions tabulated in Tables C1 and C2 of

Appendix C were computed by following the present procedure.

γ2

Ppit

P0

Ppit

P0

T∞T0

T∞T0

P∞P0

P∞P0

γ－12

T∞T0

γγ－1γ－1

2P∞P0

γγ－1γ

2γ M2－(γ－1)

γγ－1(γ＋ 1)M2

(γ－1)M2＋2

Ppit

P0

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 17

Appendix C

Tables of the operating and free-stream conditions

JAXA Research and Development Report JAXA-RR-04-035E18

.oNnuR P0 )aPM( T0 )K( M∞ ρ ∞ m/g( 3) p∞ )aP( T∞ )K( U∞ )s/m( eR (× 01 5) α )ged( skrameR0641 630.4 3.0401 56.9 22.7 1.511 5.55 6.1441 68.2 01- → 811641 630.4 1.0401 56.9 32.7 2.511 5.55 4.1441 68.2 4 → 232641 515.2 6.499 95.9 88.4 0.57 5.35 8.5041 79.1 01- → 813641 515.2 9.999 95.9 58.4 9.47 8.35 8.9041 59.1 4 → 235641 000.1 6.759 64.9 61.2 7.23 6.25 6.5731 78.0 01- → 616641 820.4 4.4401 56.9 71.7 8.411 8.55 7.4441 48.2 01- → 817641 415.2 2.0201 95.9 37.4 6.47 0.55 4.5241 88.1 01- → 818641 000.1 8.759 64.9 61.2 6.23 6.25 8.5731 78.0 6 → 239641 100.1 9.359 64.9 71.2 7.23 4.25 7.2731 88.0 03,02,01,0,01- &0741 100.1 8.839 64.9 22.2 8.23 5.15 9.0631 19.0 01- → 611741 999.0 9.439 64.9 22.2 7.23 3.15 8.7531 19.0 6 → 232741 999.0 7.249 64.9 02.2 7.23 7.15 0.4631 09.0 03,02,01,0,01-3741 000.1 5.449 64.9 02.2 7.23 8.15 4.5631 09.0 01- → 615741 650.6 5.8601 96.9 3.01 4.861 8.65 8.3641 50.4 4 → 237741 850.6 5.4401 96.9 6.01 2.961 4.55 8.5441 42.4 01- → 818741 550.6 9.1401 96.9 7.01 2.961 3.55 8.3441 62.4 4 → 230841 215.2 6.2001 95.9 38.4 8.47 9.35 9.1141 49.1 4 → 231841 799.0 9.069 64.9 51.2 5.23 8.25 2.8731 68.0 01- → 612841 820.4 9.2201 56.9 73.7 3.511 5.45 4.8241 69.2 4 → 233841 799.0 3.749 64.9 81.2 6.23 0.25 6.7631 98.0 01- → 616841 250.6 3.3701 96.9 3.01 1.861 1.75 4.7641 10.4 01- → 817841 050.6 3.5701 96.9 2.01 0.861 2.75 9.8641 00.4 43 → 6 peewsevitageN8841 150.6 4.9601 96.9 3.01 2.861 8.65 5.4641 40.4 02 → 8- peewsevitageN

)mm001(ledomehtforetemaidydobretnecehtnodesabsirebmunsdlonyeR

Pitch Pause

&Pitch Pause

.oNnuR P0 )aPM( T0 )K( M∞ ρ ∞ m/g( 3) p∞ )aP( T∞ )K( U∞ )s/m( eR (× 01 5) α )ged( φ )ged( L skrameR2231 440.6 8.8601 96.9 3.01 1.861 8.65 1.4641 40.4 0 0 1L3231 415.2 5.1201 95.9 27.4 6.47 1.55 3.6241 78.1 0 0 1L4231 220.4 9.4101 56.9 34.7 2.511 1.45 3.2241 00.3 0 0 1L9231 715.2 5.2301 95.9 66.4 6.47 7.55 7.4341 38.1 0 0 2L0331 515.2 4.7201 95.9 96.4 6.47 4.55 8.0341 58.1 0 0 3L1331 315.2 6.899 95.9 58.4 8.47 7.35 9.8041 69.1 0 081 1L2331 905.2 1.8001 95.9 97.4 6.47 3.45 1.6141 29.1 0 09 1L3331 905.2 2.2001 95.9 38.4 7.47 9.35 6.1141 49.1 0 072 1L4331 905.2 3.489 95.9 49.4 9.47 9.25 8.7931 10.2 0 0 1L5331 315.2 9.599 95.9 78.4 9.47 6.35 8.6041 79.1 51 0 1L6331 120.4 7.0401 56.9 91.7 7.411 6.55 9.1441 58.2 51 0 1L7331 040.6 8.9401 96.9 5.01 5.861 7.55 8.9441 91.4 51 0 1L8331 115.2 9.999 95.9 58.4 8.47 8.35 8.9041 59.1 51 081 1L9331 215.2 8.5201 95.9 96.4 5.47 3.55 6.9241 58.1 0 0 1L0431 905.2 0.699 95.9 68.4 8.47 6.35 9.6041 69.1 0 0 1L1431 805.2 8.2001 95.9 28.4 7.47 0.45 1.2141 49.1 0 081 1L2431 415.2 1.8001 95.9 08.4 8.47 3.45 1.6141 29.1 51 0 1L3431 215.2 4.689 95.9 39.4 0.57 0.35 4.9931 00.2 0 0 1L6431 899.0 9.919 64.9 72.2 8.23 4.05 0.6431 49.0 0 0 1L7431 415.2 3.5101 95.9 67.4 7.47 7.45 6.1241 98.1 51 0 1L8431 799.0 2.149 64.9 02.2 6.23 7.15 8.2631 09.0 0 0 1L9431 805.2 1.189 95.9 59.4 9.47 7.25 4.5931 20.2 0 0 1L aremaCRI0531 699.0 6.439 64.9 22.2 6.23 3.15 6.7531 19.0 0 0 1L aremaCRI1531 020.4 9.9401 56.9 11.7 5.411 1.65 8.8441 08.2 0 0 1L aremaCRI2531 440.6 3.8001 96.9 1.11 9.961 3.35 3.8141 45.4 0 0 1L aremaCRI3531 015.2 6.7001 95.9 08.4 7.47 2.45 7.5141 29.1 0 0 1L aremaCRI4531 215.2 6.999 95.9 58.4 8.47 8.35 6.9041 59.1 51 0 1L aremaCRI

)mm001(ledomehtforetemaidydobretnecehtnodesabsirebmunsdlonyeRmm059:3L,mm072:2L,tixeelzzonehtmorfmm005:1L

Table C1 Summery of tunnel/free-stream conditions for the force test.

Table C2 Summery of tunnel/free-stream conditions for the heat test.

Appendix D

Tables of the force/heat experimental data

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 19Ta

ble

D1

Sum

mar

y of

for

ce te

st r

esul

ts.

0641nu

R1641

nuR

2641nu

Rα

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FA

CN

Cm

α

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

352.0474.0-

916.000.8-

231.0-272.0

885.079.4

642.0764.0-

626.029.7-

702.0504.0-

906.010.7-

861.0-333.0

695.079.5

302.0993.0-

916.039.6-

071.0143.0-

306.000.6-

602.0-893.0

406.069.6

071.0633.0-

116.039.5-

331.0672.0-

595.099.4-

252.0-864.0

416.079.7

721.0272.0-

306.039.4-

790.0512.0-

785.000.4-

803.0-445.0

626.089.8

590.0212.0-

695.029.3-

660.0951.0-

385.099.2-

773.0-426.0

636.089.9

360.0551.0-

095.019.2-

930.0401.0-

875.099.1-

564.0-517.0

056.089.01

140.0101.0-

785.019.1-

810.0150.0-

675.089.0-

775.0-518.0

666.099.11

510.0940.0-

585.029.0-

000.0100.0-

675.010.0

717.0-829.0

386.099.21

200.0-100.0

585.060.0

120.0-250.0

675.010.1

888.0-550.1

207.099.31

120.0-250.0

685.070.1

240.0-301.0

875.010.2

690.1-591.1

427.099.41

140.0-501.0

985.070.2

960.0-851.0

385.020.3

643.1-553.1

057.000.61

070.0-161.0

495.080.3

201.0-712.0

095.020.4

226.1-825.1

087.000.71

001.0-812.0

006.080.4

731.0-872.0

795.030.5

019.1-907.1

218.000.81

831.0-972.0

706.080.5

371.0-143.0

506.030.6

131.2-378.1

938.010.91

571.0-343.0

516.090.6

212.0-704.0

416.030.7

103.2-520.2

568.010.02

412.0-904.0

526.080.7

952.0-774.0

326.040.8

534.2-561.2

788.010.12

162.0-084.0

436.090.8

313.0-355.0

236.040.9

245.2-892.2

609.020.22

123.0-755.0

546.001.9

583.0-336.0

246.040.01

746.2-134.2

629.020.32

193.0-936.0

756.001.01

674.0-527.0

556.050.11

057.2-265.2

449.030.42

264.0-117.0

766.009.01

985.0-628.0

176.050.21

458.2-496.2

169.040.52

475.0-118.0

286.009.11

237.0-149.0

986.050.31

569.2-138.2

779.040.62

907.0-229.0

896.019.21

909.0-960.1

807.050.41

770.3-869.2

599.040.72

878.0-740.1

717.019.31

221.1-312.1

137.060.51

591.3-901.3

110.150.82

380.1-881.1

937.019.41

173.1-273.1

657.060.61

113.3-842.3

620.160.92

723.1-443.1

467.029.51

266.1-355.1

887.070.71

824.3-583.3

040.170.03

906.1-025.1

597.029.61

055.3-925.3

550.170.13

029.1-027.1

138.060.81

3641nu

R5641

nuR

6641nu

Rα

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FA

CN

Cm

α

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FA

CN

Cm

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CN

Cm

731.0-082.0

406.090.5

752.0184.0-

576.070.8-

842.0864.0-

316.089.7-

371.0-343.0

316.080.6

312.0414.0-

666.070.7-

302.0104.0-

406.089.6-

312.0-904.0

126.080.7

661.0543.0-

456.060.6-

661.0533.0-

795.089.5-

852.0-874.0

136.090.8

131.0772.0-

056.060.5-

721.0272.0-

885.089.4-

813.0-555.0

346.001.9

390.0512.0-

246.060.4-

490.0212.0-

185.079.3-

783.0-636.0

556.090.01

360.0551.0-

536.050.3-

360.0651.0-

675.069.2-

084.0-927.0

866.001.11

830.0201.0-

336.050.2-

930.0201.0-

475.069.1-

095.0-828.0

386.001.21

220.0150.0-

136.050.1-

710.0150.0-

175.069.0-

107.0-619.0

796.009.21

000.0100.0

826.050.0-

100.0-100.0-

075.030.0

868.0-240.1

717.009.31

510.0-250.0

136.049.0

020.0-150.0

275.040.1

570.1-481.1

937.009.41

140.0-501.0

136.049.1

140.0-201.0

675.030.2

023.1-243.1

567.019.51

560.0-161.0

836.059.2

070.0-751.0

085.040.3

795.1-715.1

597.019.61

101.0-022.0

346.059.3

101.0-612.0

785.040.4

688.1-796.1

628.019.71

231.0-182.0

256.069.4

731.0-672.0

395.050.5

421.2-668.1

658.029.81

861.0-643.0

066.069.5

371.0-833.0

106.060.6

003.2-020.2

288.029.91

902.0-314.0

076.059.6

212.0-504.0

016.050.7

634.2-061.2

409.029.02

062.0-584.0

286.069.7

852.0-574.0

026.060.8

255.2-592.2

529.039.12

513.0-165.0

396.069.8

713.0-155.0

036.070.9

656.2-924.2

549.049.22

093.0-646.0

707.079.9

783.0-336.0

246.070.01

067.2-065.2

369.049.32

774.0-837.0

127.079.01

974.0-427.0

456.080.11

768.2-696.2

289.049.42

485.0-938.0

737.079.11

095.0-528.0

076.080.21

779.2-238.2

899.059.52

727.0-259.0

457.079.21

437.0-939.0

686.080.31

980.3-869.2

510.159.62

009.0-670.1

277.079.31

809.0-760.1

707.070.41

602.3-701.3

130.159.72

701.1-122.1

697.089.41

921.1-512.1

037.080.51

913.3-342.3

640.169.82

353.1-973.1

528.079.51

773.1-273.1

657.090.61

144.3-783.3

260.179.92

066.1-945.1

787.090.71

065.3-825.3

770.179.03

919.1-717.1

618.040.81

7641nu

R8641

nuR

9641nu

Rα

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

152.0074.0-

036.020.8-

662.0-884.0

086.070.8

383.0346.0-

507.040.01-

502.0304.0-

126.020.7-

623.0-665.0

196.070.9

400.0-300.0-

236.010.0-

861.0833.0-

316.010.6-

693.0-256.0

407.080.01

583.0-046.0

917.000.01

921.0572.0-

506.010.5-

784.0-147.0

717.080.11

163.2-760.2

759.050.02

490.0512.0-

895.010.4-

306.0-848.0

437.080.21

645.3-764.3

351.121.03

260.0851.0-

295.000.3-

747.0-169.0

257.080.31

040.0401.0-

985.000.2-

619.0-880.1

477.080.41

810.0250.0-

685.099.0-

421.1-432.1

008.090.51

100.0-100.0-

785.000.0

673.1-293.1

528.090.61

020.0-050.0

885.000.1

995.1-635.1

948.009.61

040.0-201.0

095.000.2

259.1-857.1

098.001.81

760.0-851.0

595.010.3

141.2-198.1

119.019.81

001.0-512.0

206.010.4

833.2-350.2

049.019.91

531.0-672.0

016.020.5

294.2-002.2

469.009.02

271.0-933.0

816.020.6

716.2-043.2

789.029.12

212.0-604.0

726.010.7

327.2-574.2

800.129.22

852.0-774.0

736.020.8

148.2-616.2

920.129.32

813.0-355.0

846.030.9

949.2-457.2

840.129.42

783.0-636.0

066.030.01

260.3-198.2

560.139.52

874.0-727.0

376.040.11

771.3-430.3

680.139.62

785.0-628.0

786.030.21

003.3-871.3

401.139.72

927.0-049.0

407.040.31

714.3-913.3

021.149.82

509.0-860.1

427.040.41

535.3-064.3

631.149.92

811.1-312.1

747.050.51

746.3-006.3

051.159.03

763.1-473.1

477.050.61

356.1-255.1

508.060.71

339.1-037.1

738.050.81

0741nu

R1741

nuR

2741nu

Rα

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

062.0294.0-

376.030.8-

652.0-564.0

376.079.7

483.0256.0-

886.040.01-

721.0092.0-

246.020.5-

474.0-617.0

017.089.01

163.2-950.2

739.050.02

990.0822.0-

936.030.4-

385.0-718.0

427.089.11

745.3-064.3

031.121.03

160.0171.0-

926.020.3-

027.0-039.0

247.089.21

640.0811.0-

926.020.2-

588.0-450.1

367.089.31

810.0560.0-

626.020.1-

501.1-991.1

687.089.41

800.0510.0-

626.020.0-

933.1-653.1

418.099.51

710.0-830.0

526.079.0

026.1-335.1

548.099.61

630.0-290.0

926.079.1

019.1-817.1

878.000.81

560.0-841.0

536.089.2

561.2-198.1

609.000.91

490.0-802.0

446.089.3

543.2-840.2

539.000.02

231.0-762.0

846.089.4

894.2-491.2

959.000.12

861.0-233.0

956.099.5

516.2-233.2

289.010.22

212.0-104.0

766.089.6

337.2-174.2

300.110.32

552.0-374.0

086.099.7

838.2-606.2

320.120.42

323.0-255.0

986.099.8

549.2-047.2

040.120.52

683.0-436.0

207.000.01

950.3-188.2

160.120.62

184.0-527.0

617.000.11

181.3-320.3

770.120.72

585.0-328.0

037.010.21

782.3-261.3

490.130.82

437.0-739.0

547.010.31

224.3-803.3

211.130.92

109.0-560.1

967.010.41

825.3-744.3

921.140.03

411.1-212.1

497.010.51

656.3-495.3

541.150.13

243.1-563.1

128.010.61

JAXA Research and Development Report JAXA-RR-04-035E20

2841nu

R3841

nuR

6841nu

Rα

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

321.0-262.0

385.069.4

062.0184.0-

876.050.8-

152.0664.0-

006.010.8-

461.0-523.0

985.000.6

312.0514.0-

866.050.7-

502.0993.0-

195.010.7-

302.0-193.0

895.010.7

471.0253.0-

166.040.6-

761.0433.0-

485.030.6-

742.0-954.0

806.010.8

821.0582.0-

156.040.5-

131.0272.0-

675.020.5-

203.0-435.0

026.059.8

990.0522.0-

446.040.4-

690.0212.0-

075.020.4-

273.0-516.0

036.000.01

460.0561.0-

836.030.3-

560.0651.0-

465.010.3-

654.0-207.0

446.079.01

630.0801.0-

436.030.2-

140.0201.0-

265.080.2-

265.0-008.0

856.069.11

710.0650.0-

136.030.1-

020.0150.0-

955.050.1-

696.0-019.0

376.000.31

100.0-600.0-

236.030.0-

000.0100.0-

755.010.0

668.0-530.1

396.069.31

220.0-840.0

036.069.0

910.0-940.0

165.049.0

370.1-671.1

417.010.51

730.0-101.0

536.069.1

830.0-001.0

565.039.1

023.1-433.1

247.099.51

460.0-551.0

936.079.2

560.0-551.0

965.069.2

795.1-805.1

277.089.61

790.0-512.0

846.079.3

790.0-212.0

575.029.3

788.1-196.1

508.069.71

831.0-772.0

356.089.4

231.0-172.0

285.049.4

111.2-558.1

338.079.81

671.0-343.0

366.089.5

471.0-643.0

195.021.6

982.2-900.2

858.000.02

612.0-014.0

176.089.6

502.0-793.0

795.069.6

714.2-641.2

978.099.02

062.0-284.0

286.089.7

952.0-184.0

016.070.8

235.2-182.2

109.079.12

423.0-065.0

396.099.8

603.0-245.0

516.030.9

136.2-114.2

919.079.22

193.0-346.0

707.099.9

573.0-326.0

726.089.9

537.2-145.2

639.020.42

884.0-837.0

127.099.01

364.0-317.0

046.089.01

738.2-476.2

459.000.52

695.0-838.0

637.000.21

275.0-118.0

556.049.11

949.2-018.2

079.010.62

737.0-159.0

657.000.31

217.0-429.0

376.039.21

260.3-749.2

689.040.72

809.0-180.1

677.000.41

388.0-940.1

196.019.31

871.3-480.3

100.150.82

711.1-722.1

408.000.51

690.1-491.1

517.029.41

492.3-522.3

810.140.92

543.1-183.1

828.010.61

543.1-453.1

247.019.51

114.3-363.3

230.180.03

726.1-825.1

177.049.61

925.3-305.3

740.160.13

439.1-627.1

508.050.81

7841nu

R8841

nuR

α

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

571.0-153.0

595.022.6

712.0614.0-

895.002.7-

491.0-583.0

895.018.6

771.0053.0-

095.091.6-

732.0-254.0

606.018.7

141.0782.0-

285.091.5-

882.0-625.0

616.028.8

701.0622.0-

475.081.4-

353.0-406.0

626.038.9

570.0961.0-

965.081.3-

434.0-096.0

736.038.01

640.0411.0-

465.081.2-

735.0-687.0

156.038.11

620.0360.0-

265.071.1-

666.0-398.0

666.048.21

500.0210.0-

955.071.0-

928.0-510.1

486.038.31

210.0-830.0

265.028.0

720.1-251.1

507.058.41

330.0-980.0

365.038.1

462.1-403.1

927.048.51

850.0-241.0

665.038.2

335.1-374.1

757.058.61

880.0-891.0

275.048.3

228.1-356.1

887.058.71

021.0-652.0

975.048.4

160.2-228.1

718.058.81

751.0-813.0

685.058.5

142.2-479.1

148.068.91

491.0-283.0

395.058.6

873.2-311.2

368.078.02

532.0-944.0

206.068.7

294.2-542.2

288.078.12

882.0-325.0

116.068.8

095.2-473.2

009.088.22

253.0-995.0

126.078.9

196.2-205.2

719.098.32

334.0-686.0

336.078.01

197.2-926.2

239.098.42

535.0-187.0

646.088.11

109.2-767.2

059.019.52

466.0-788.0

166.088.21

010.3-898.2

469.019.62

228.0-500.1

776.088.31

221.3-330.3

879.009.72

420.1-541.1

007.098.41

532.3-661.3

499.029.82

952.1-592.1

427.088.51

353.3-303.3

700.139.92

725.1-264.1

157.098.61

874.3-344.3

020.189.03

808.1-746.1

387.098.71

Tabl

e D

1Su

mm

ary

of f

orce

test

res

ults

. (co

ntin

ued)

3741nu

R5741

nuR

7741nu

Rα

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

362.0294.0-

176.080.8-

421.0-662.0

085.098.4

832.0354.0-

695.039.7-

612.0224.0-

466.070.7-

561.0-823.0

685.019.5

791.0683.0-

985.049.6-

371.0653.0-

556.060.6-

302.0-293.0

495.019.6

161.0323.0-

185.039.5-

531.0092.0-

946.060.5-

642.0-064.0

306.029.7

621.0262.0-

475.039.4-

190.0622.0-

146.060.4-

103.0-435.0

216.029.8

290.0302.0-

865.039.3-

560.0171.0-

536.060.3-

863.0-316.0

326.039.9

160.0741.0-

265.029.2-

140.0711.0-

726.050.2-

554.0-207.0

736.039.01

630.0490.0-

855.019.1-

320.0460.0-

726.050.1-

265.0-008.0

156.039.11

610.0340.0-

755.019.0-

000.0110.0-

626.050.0-

007.0-219.0

866.039.21

200.0-600.0

655.080.0

710.0-140.0

726.059.0

078.0-730.1

686.039.31

220.0-650.0

855.090.1

730.0-490.0

036.049.1

080.1-971.1

907.049.41

340.0-701.0

165.090.2

160.0-051.0

636.059.2

423.1-533.1

437.049.51

070.0-261.0

665.001.3

390.0-902.0

446.059.3

106.1-705.1

467.059.61

201.0-912.0

275.001.4

731.0-172.0

056.069.4

988.1-096.1

697.059.71

921.0-762.0

875.019.4

171.0-533.0

956.069.5

011.2-458.1

428.069.81

661.0-923.0

585.019.5

802.0-104.0

566.069.6

772.2-300.2

948.069.91

302.0-393.0

395.019.6

852.0-274.0

476.079.7

904.2-241.2

078.059.02

842.0-264.0

206.029.7

113.0-945.0

786.079.8

515.2-372.2

988.079.12

203.0-735.0

216.029.8

983.0-136.0

896.079.9

616.2-404.2

809.089.22

173.0-716.0

326.039.9

674.0-527.0

417.089.01

717.2-435.2

629.099.32

854.0-607.0

636.049.01

495.0-828.0

237.089.11

228.2-766.2

249.099.42

865.0-408.0

946.049.11

627.0-939.0

847.089.21

339.2-408.2

069.099.52

307.0-519.0

666.049.21

198.0-460.1

967.089.31

640.3-149.2

679.000.72

878.0-240.1

586.069.31

501.1-212.1

497.099.41

061.3-080.3

199.010.82

480.1-281.1

707.069.41

333.1-363.1

228.079.51

872.3-812.3

700.120.92

333.1-043.1

337.089.51

593.3-753.3

120.130.03

906.1-315.1

267.089.61

715.3-105.3

630.130.13

398.1-396.1

497.099.71

8741nu

R0841

nuR

1841nu

Rα

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

α

)ged(C

FA

CN

Cm

531.0-872.0

475.080.5

231.0-772.0

406.010.5

552.0574.0-

876.039.7-

171.0-833.0

285.070.6

861.0-733.0

216.010.6

002.0993.0-

076.049.6-

902.0-304.0

295.070.7

112.0-104.0

026.010.7

851.0233.0-

956.039.5-

452.0-274.0

106.080.8

352.0-274.0

136.010.8

821.0072.0-

156.039.4-

013.0-745.0

016.090.9

613.0-155.0

246.020.9

980.0012.0-

446.039.3-

083.0-726.0

226.080.01

083.0-136.0

456.020.01

060.0151.0-

836.029.2-

864.0-517.0

436.090.11

274.0-127.0

766.030.11

230.0990.0-

536.029.1-

285.0-618.0

056.090.21

185.0-228.0

486.030.21

210.0440.0-

436.029.0-

327.0-929.0

666.090.31

427.0-639.0

007.030.31

700.0-600.0

036.070.0

898.0-650.1

586.090.41

198.0-260.1

227.030.41

810.0-950.0

436.080.1

211.1-002.1

907.001.51

011.1-802.1

547.030.51

740.0-411.0

636.070.2

013.1-623.1

927.019.51

553.1-663.1

077.040.61

960.0-961.0

146.080.3

685.1-994.1

957.019.61

246.1-545.1

108.040.71

401.0-822.0

846.080.4

078.1-776.1

097.019.71

529.1-527.1

338.050.81

931.0-192.0

556.090.5

290.2-148.1

818.029.81

351.2-398.1

168.050.91

671.0-653.0

666.090.6

162.2-989.1

248.039.91

523.2-740.2

988.060.02

712.0-424.0

476.090.7

093.2-721.2

368.029.02

564.2-881.2

119.050.12

662.0-594.0

286.090.8

005.2-952.2

388.049.12

175.2-913.2

039.070.22

723.0-475.0

696.001.9

895.2-883.2

109.049.22

876.2-354.2

159.070.32

983.0-046.0

307.009.9

996.2-815.2

919.059.32

187.2-685.2

869.070.42

574.0-237.0

127.009.01

508.2-056.2

639.069.42

298.2-127.2

689.080.52

195.0-238.0

437.009.11

419.2-687.2

359.069.52

400.3-368.2

500.180.62

817.0-549.0

557.019.21

420.3-129.2

969.079.62

811.3-699.2

910.180.72

898.0-770.1

277.019.31

241.3-160.3

589.079.72

332.3-831.3

730.190.82

590.1-612.1

897.019.41

652.3-791.3

999.089.82

353.3-772.3

250.101.92

043.1-773.1

728.059.51

073.3-333.3

310.199.92

944.3-293.3

660.119.92

194.3-574.3

820.100.13

765.3-335.3

180.119.03

Comparative Force/Heat Flux Measurements between JAXA Hypersonic Test Facilities Using Standard Model HB-2 (Part 1: 1.27 m Hypersonic Wind Tunnel Results) 21

.oNrosneS m/Wk(xulFtaeH 2)2231 3231 4231 7231 8231 1331 2331 3331 4331 5331 6331 7331 8331 9331

1 27.19 12.65 30.07 09.55 67.55 59.35 22.55 31.45 15.35 95.92 51.93 63.84 63.48 49.552 22.751 63.69 48.811 19.49 04.59 79.29 27.49 57.29 25.19 98.16 75.18 36.001 26.021 39.593 62.412 17.231 88.361 30.131 74.131 02.921 11.131 14.821 77.621 31.401 42.631 25.761 09.041 16.2314 80.232 69.341 88.771 15.241 23.341 88.041 23.241 61.041 26.731 33.231 51.371 44.212 85.231 32.4415 63.902 14.921 62.061 02.821 95.821 50.721 78.721 45.621 67.321 63.931 32.281 18.222 98.001 44.9216 10.551 22.59 81.811 74.49 87.49 97.39 41.49 66.39 99.09 33.121 08.851 04.491 86.16 71.597 05.09 44.55 90.96 51.55 14.55 41.55 91.55 26.45 39.25 22.58 22.211 19.731 77.92 86.558 55.25 09.13 53.04 19.13 01.23 41.23 47.13 28.13 95.03 39.45 30.27 51.98 71.51 03.239 64.24 37.52 12.23 65.52 27.52 68.52 05.52 37.52 29.42 29.84 44.46 87.97 30.11 39.5211 72.31 61.8 22.01 22.8 22.8 05.8 03.8 62.8 38.7 00.91 90.52 11.13 95.2 24.821 55.7 19.4 00.6 65.4 80.5 28.4 28.4 38.4 84.4 62.11 31.51 00.91 41.1 18.431 81.5 41.3 97.3 11.3 80.3 74.3 31.3 03.3 20.3 32.11 14.51 32.91 20.1 24.341 77.3 42.2 00.3 51.2 75.2 55.2 02.2 74.2 61.2 25.11 74.51 91.91 24.2 41.251 84.3 26.1 36.2 61.2 61.2 92.2 59.1 93.2 10.2 14.11 61.51 37.81 10.3 71.261 59.2 96.1 33.2 59.1 41.2 71.2 88.1 91.2 41.2 40.11 50.51 86.81 13.3 38.171 35.2 45.1 41.2 49.1 89.1 10.2 47.1 60.2 85.1 91.21 45.61 15.02 24.4 08.181 59.2 88.1 75.2 21.2 40.2 21.2 58.1 39.1 79.1 16.61 12.22 91.82 18.5 80.291 75.4 87.2 25.3 17.2 87.2 09.2 36.2 38.2 88.2 84.22 42.03 22.73 27.7 49.202 81.5 59.2 88.3 60.3 87.2 14.3 32.3 07.3 91.3 36.42 17.23 52.04 87.7 72.312 82.5 13.3 81.4 11.3 52.3 18.3 12.3 26.3 23.3 10.72 21.63 75.44 12.7 45.322 05.5 51.3 81.4 13.3 83.3 38.3 53.3 54.3 22.3 79.61 27.22 10.82 29.0 03.332 62.5 62.3 01.4 32.3 64.3 97.3 66.3 03.3 71.3 99.5 14.7 60.9 79.5 43.342 17.5 75.3 93.4 65.3 46.3 83.3 47.3 51.3 53.3 39.0 10.1 02.1 62.71 86.352 52.6 68.3 30.5 55.3 58.3 62.3 66.3 44.3 59.3 48.6 27.9 90.21 98.62 29.362 93.6 67.3 30.5 19.3 08.3 73.3 18.3 73.3 63.3 29.0 00.1 40.1 10.71 29.372 68.5 45.3 27.4 15.3 87.3 05.3 12.3 65.3 73.3 31.6 96.7 30.9 79.5 46.382 86.5 04.3 24.4 04.3 77.3 04.3 32.3 47.3 62.3 80.71 65.22 78.72 07.0 53.3

.oNrosneS m/Wk(xulFtaeH 2)0431 1431 2431 3431 6431 7431 8431 9431 0531 1531 2531 3531 4531

1 26.45 30.45 75.92 59.35 59.13 22.03 97.232 59.29 78.29 38.16 98.19 29.45 59.26 90.653 76.821 48.821 30.401 80.721 90.67 33.601 79.774 08.931 06.041 26.231 29.731 50.38 32.531 19.485 06.521 67.621 44.931 71.421 85.47 80.241 80.676 07.29 76.39 58.121 03.19 06.45 39.321 77.557 31.45 99.45 25.58 22.35 75.13 10.78 23.238 70.13 13.23 48.45 87.03 90.81 43.65 57.819 02.52 09.52 31.94 28.42 05.41 99.94 00.5111 88.7 05.8 61.91 68.7 67.4 28.91 79.421 35.4 78.4 65.11 48.4 56.2 76.11 76.231 40.3 77.3 27.11 60.3 04.1 56.11 58.1 41.3 39.1 43.4 31.5 62.3 19.1141 55.2 46.2 55.11 33.2 24.1 38.11 55.1 03.2 04.1 50.3 56.3 16.2 96.1151 32.2 82.2 32.11 52.2 11.1 56.11 74.1 10.2 54.1 86.2 62.3 50.2 05.1161 79.1 11.2 31.11 29.1 40.1 04.11 53.1 50.2 82.1 56.2 31.3 89.1 73.1171 68.1 20.2 92.21 95.1 77.0 67.21 12.1 40.2 70.1 45.2 57.2 48.1 76.2181 57.1 61.2 98.61 50.2 91.1 11.71 14.1 30.2 22.1 94.2 77.2 89.1 70.7191 26.2 30.3 56.22 76.2 36.1 09.22 47.1 18.2 78.1 57.3 03.4 56.2 88.2202 19.2 42.3 55.42 03.3 47.1 30.52 85.1 23.3 88.1 04.4 96.4 51.3 08.4212 33.3 39.3 51.72 53.3 88.1 66.72 58.1 64.3 57.1 44.4 52.5 61.3 41.7222 90.3 74.3 92.71 22.3 97.1 94.71 78.1 32.3 11.2 32.4 02.5 12.3 41.7132 84.3 64.3 20.6 83.3 80.2 58.5 80.2 33.3 61.2 47.4 89.4 13.3 38.542 33.3 35.3 00.1 14.3 28.1 61.1 50.2 93.3 25.2 15.4 67.5 26.3 78.052 06.3 24.3 29.6 07.3 91.2 01.7 01.2 66.3 15.2 72.5 99.5 78.3 22.762 36.3 95.3 96.0 35.3 12.2 91.1 38.1 57.3 62.2 87.4 56.5 74.3 29.072 64.3 15.3 58.5 24.3 21.2 20.6 91.2 95.3 70.2 20.5 14.5 97.3 50.682 24.3 83.3 91.71 92.3 96.1 34.71 40.2 54.3 11.2 45.4 80.5 11.3 63.71

Table D2 Summary of heat test results.

JAXA Research and Development Report

(JAXA-RR-04-035E)

Date of Issue: March 31, 2005

Edited and Published by:Japan Aerospace Explortion Agency7-44-1 Jindaiji-higashimachi, Chofu-shi,Tokyo 182-8522 Japan

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�2005 JAXA, All Right Reserved

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