Hypersonic aerothermodynamic and scramjet research using high enthalpy shock tunnel

Itoh, Katsuhiro; Ueda, Shuichi; Tanno, Hideyuki; Komuro, Tomoyuki; Sato, K.

doi:10.1007/s00193-002-0147-0

Cited by 86 publications

(22 citation statements)

References 5 publications

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“…A spacecraft re-entry simulation is conducted using a shock tube and a wind tunnel. He test times for some of the equipment are extremely short [6]. Therefore we cannot use standard TSP because it does not have a sufficiently short response time.…”

Section: Introductionmentioning

confidence: 99%

Development of Temperature-Sensitive Paint with High Performance and Responsibility for Aerodynamic Heating Measurement

Horagiri

Nagai

2014

52nd Aerospace Sciences Meeting

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Heat flux measurement data are indispensable in the design of a heat protection system for the atmospheric re-entry of a spacecraft. One of the instrumentation techniques used to obtain heat flux data is Temperature-Sensitive Paint (TSP). Shock tubes and shock tunnel are used to simulate re-entry, but the duration of a shock tunnel test is extremely short. In order to apply TSP to high speed phenomena, a new TSP with a large luminescence and high responsivity was developed. In a luminescence test, the luminescent intensity peaked when the mixing ratio was set at 50%, and the intensity was about twice the original one. In a response time evaluation test, the response time was improved by increasing the mixing ratio. When the calcium silicate mixing ratio was set at 50%, the improved TSP achieved a response with a time constant of 0.205ms. We applied this TSP to a high-enthalpy wind tunnel and could obtain the heat flux data. Nomenclature q = heat flux, W m -2 q norm = normalized with a maximum heat flux c = specific heat, J kg -1 K -1 I = luminescent intensity t = time, ms a = thermal diffusivity, m 2 s -1 k = thermal conductivity, J s -1 m -1 K -1 ρ = density, kg m -3 L = film thickness, m A n = calibration coefficient T = temperature, K τ = time constant, ms θ = temperature change from initial temperature, K W = error function subscript: ref = reference condition max = maximum value b = base p = polymer 1 Master's course student,

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Section: Introductionmentioning

confidence: 99%

Development of Temperature-Sensitive Paint with High Performance and Responsibility for Aerodynamic Heating Measurement

Horagiri

Nagai

2014

52nd Aerospace Sciences Meeting

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“…(center of pressure), so that the total pitch moment produces a higher nose-up attitude than under perfect-gas conditions. The high-enthalpy shock tunnel HIEST 4 (Fig. 1) was originally built to study such hightemperature real-gas effects on reentry vehicle aerodynamics.…”

Section: Introductionmentioning

confidence: 99%

Free-flight measurement technique in the free-piston high-enthalpy shock tunnel

Tanno

Komuro

Sato

et al. 2014

Review of Scientific Instruments

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Articles you may be interested inOn-road and wind-tunnel measurement of motorcycle helmet noise J. Acoust. Soc. Am. 134, 2004Am. 134, (2013 A novel multi-component force-measurement technique has been developed and implemented at the impulse facility JAXA-HIEST, in which the test model is completely unrestrained during the test and thus experiences free-flight conditions for a period on the order of milliseconds. Advantages over conventional free-flight techniques include the complete absence of aerodynamic interference from a model support system and less variation in model position and attitude during the test itself. A miniature on-board data recorder, which was a key technology for this technique, was also developed in order to acquire and store the measured data. The technique was demonstrated in a HIEST windtunnel test campaign in which three-component aerodynamic force measurement was performed on a blunted cone of length 316 mm, total mass 19.75 kg, and moment of inertia 0.152 kgm 2 . During the test campaign, axial force, normal forces, and pitching moment coefficients were obtained at angles of attack from 14 • to 32 • under two conditions: H 0 = 4 MJ/kg, P 0 = 14 MPa; and H 0 = 16 MJ/kg, P 0 = 16 MPa. For the first, low-enthalpy condition, the test flow was considered a perfect gas; measurements were thus directly compared with those obtained in a conventional blow-down wind tunnel (JAXA-HWT2) to evaluate the accuracy of the technique. The second test condition was a high-enthalpy condition in which 85% of the oxygen molecules were expected to be dissociated; high-temperature real-gas effects were therefore evaluated by comparison with results obtained in perfect-gas conditions. The precision of the present measurements was evaluated through an uncertainty analysis, which showed the aerodynamic coefficients in the HIEST low enthalpy test agreeing well with those of JAXA-HWT2. The pitching-moment coefficient, however, showed significant differences between low-and high-enthalpy tests. These differences are thought to result from hightemperature real-gas effects. © 2014 AIP Publishing LLC. [http://dx

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“…These techniques include using very light models with stiff balances and the application of vibration compensation methods via the use of accelerometers [6][7][8]. Other techniques include using free-flying or weakly constrained models [9][10][11][12][13][14][15][16] and either directly measuring the accelerations or deriving these from displacement profiles in order to infer the applied aerodynamic forces.…”

Section: Introductionmentioning

confidence: 99%

Design and implementation of an internal stress wave force balance in a shock tunnel

2010

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A stress wave internal force balance for the High Enthalpy Shock Tunnel Göttingen (HEG) of the German Aerospace Center (DLR) to measure lift, pitching moment and drag was designed, calibrated and tested. The balance is designed to measure forces in ground based test facilities with test times in the order of milliseconds on models additionally instrumented with surface pressure and wall heat flux gauges from angles of attack of -40°to 20°. Experiments in HEG were performed on a 303 mm long, 10°half angle blunt cone at angles of attack from -20°t o 0°. The tests were conducted utilizing two different operating conditions at total specific enthalpies of 3.0 and 3.5 MJ/kg and dynamic pressures of 30 and 72 kPa. The performance of the balance was assessed by comparing the measured force and moment coefficients with computational fluid dynamics (CFD) predictions.

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Hypersonic aerothermodynamic and scramjet research using high enthalpy shock tunnel

Cited by 86 publications

References 5 publications

Development of Temperature-Sensitive Paint with High Performance and Responsibility for Aerodynamic Heating Measurement

Development of Temperature-Sensitive Paint with High Performance and Responsibility for Aerodynamic Heating Measurement

Free-flight measurement technique in the free-piston high-enthalpy shock tunnel

Design and implementation of an internal stress wave force balance in a shock tunnel

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