Hideyuki Tanno

Time-Resolved Visualization of Instability Waves in a Hypersonic Boundary Layer

LAMINAR-TURBULENT transition in hypersonic boundary layers remains a challenging subject. This is especially true of the hypervelocity regime, in which an intriguing phenomenon is the possible damping of second-mode disturbances by chemical and vibrational nonequilibrium processes. To generate flows with sufficiently high enthalpy to investigate such effects, the use of shock-tunnel facilities is necessary; furthermore, it is now generally accepted that direct measurements of the instability mechanisms active within the boundary layer, together with a characterization of the freestream disturbance environment, are required, as simple measurements of transition locations can lead to ambiguous conclusions. However, as difficult as the accurate measurement of instability waves in conventional hypersonic facilities can be, in shock tunnels it is appreciably more so. For identical unit Reynolds numbers, the higher stagnation temperature in a shock tunnel means that the dominant second-mode disturbances lie at even higher frequencies (typically hundreds of kHz or higher); moreover, because of the destructive testing environment, hot-wire techniques, a staple for instability measurements in conventional tunnels, cannot be used. Fast-response pressure transducers are an obvious alternative, but recent experiments have highlighted the challenging nature of interpreting data from mechanically sensitive sensors in the high-noise environment of a shock tunnel, especially without accompanying stability computations. Measurements with recently developed atomic-layer thermopile (ALTP) heat-flux sensors show promise, though their use has yet to be demonstrated in shocktunnel facilities.

Influence of Thermal Non-equilibrium on Nozzle Flow Condition of High Enthalpy Shock Tunnel HIEST

Axisymmetric RANS simulations using thermal and chemical non-equilibrium models are applied to a nozzle flow of a free-piston driven shock tunnel HIEST. The static pressure and Pitot pressure profiles at the nozzle exit are compared with the experimental results to examine influence of thermal non-equilibrium on the nozzle flow and validate the CFD code. As the thermal non-equilibrium model, the two-temperature model of Park and a fourtemperature model, for which the conservation of the vibrational-electronic excitation energy for N2, O2 and NO is taken into account individually, are applied to the code. The results show that the HIEST nozzle flow is close to thermal equilibrium with air as the test gas and thermal non-equilibrium with nitrogen. Good agreement between the CFD and the experimental results is obtained at the total enthalpy conditions of 3.81 MJ/kg with air and 7.99 MJ/kg with nitrogen. However, some discrepancy in the Pitot pressure profiles appears at the 10.0 and 19.0 MJ/kg conditions with air. The CFD with the 4T model does not simulate the air nozzle flow at the 3.81 MJ/kg condition. The results indicate that the 4T model underestimates the vibrational relaxation rate.

Aerodynamic force measurement on a large-scale model in a short duration test facility

A force measurement technique has been developed for large-scale aerodynamic models with a short test time. The technique is based on direct acceleration measurements, with miniature accelerometers mounted on a test model suspended by wires. Measuring acceleration at two different locations, the technique can eliminate oscillations from natural vibration of the model. The technique was used for drag force measurements on a 3m long supersonic combustor model in the HIEST free-piston driven shock tunnel. A time resolution of 350μs is guaranteed during measurements, whose resolution is enough for ms order test time in HIEST. To evaluate measurement reliability and accuracy, measured values were compared with results from a three-dimensional Navier–Stokes numerical simulation. The difference between measured values and numerical simulation values was less than 5%. We conclude that this measurement technique is sufficiently reliable for measuring aerodynamic force within test durations of 1ms.

Computation and Experiment of Nonequilibrium Nozzle Flow of Arc-heated Air

Numerical simulations of a nonequilibrium nozzle flow of are-heated air were carried out using an eight-temperature model composed of translational, N 2 -rotational, O 2 -rotational, NO-rotational, N 2 -vibrational, O 2 -vibrational, NO-vibrational, and electron temperatures. The on-axis profile of each temperature in the nozzle is shown and the thermal characteristics of the nozzle flow are discussed. Measurements of NO emission spectra were also made at wavelengths 220-265 nm at the nozzle exit to determine NO rotational temperature by a curve-fitting method. Computed rotational temperature of NO at the nozzle exit was compared with the experimental temperature to discuss the nozzle flow model introduced to the present numerical analysis. Moreover, the present computation was applied to a nozzle flow in another arcjet facility

Heat flux measurement of Apollo capsule model in the free-piston shock tunnel HIEST

This report presents the latest test results on the aeroheating of an Apollo capsule model under high-enthalpy flow conditions in the free-piston shock tunnel HIEST at JAXA Kakuda. The objective was to obtain benchmark data to validate JAXA in-house numerical simulation codes developed to predict hypervelocity flow fields such as the external flow around a re-entry vehicle. The SUS-304 stainless steel model had a maximum diameter of 250 mm and was equipped with 84 miniature co-axial thermocouples on its windward surface. Twelve thermocouples were also mounted on the leeward side of the model and were used to determine the establishment of flow around the model. Heat flux distribution around the model was measured in a high-enthalpy, high-pressure (i.e. high Reynolds number) flow at 0o and 30o angles of attack. Aeroheating characteristics were observed with a fully laminar boundary layer and with a transition boundary layer. To change the free-stream Reynolds number, stagnation pressure was varied from 14 MPa to 54 MPa and stagnation enthalpy was varied from 3.6 MJ/kg to 21 MJ/kg. The flow characteristics of the graphite nozzle throat, which was designed for high-enthalpy, high-pressure shots, are also discussed.

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

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 wind-tunnel 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: H0 = 4 MJ/kg, P0 = 14 MPa; and H0 = 16 MJ/kg, P0 = 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 high-temperature real-gas effects.

Free-flight force measurement technique in a shock tunnel

A novel force measurement technique has been developed at the impulsive facility HIEST, in which the test model is completely non-restrained for the duration of the test, so it experiences completely free-flight conditions for a period on the order of milliseconds. This technique was demonstrated with a three-component aerodynamic force measurement with a blunted cone of total length 318 mm and a total mass of 22 kg in hypervelocity test flow. A miniature modelonboard data-logger, which was a key technology for this technique, was also developed in order to store the measured data. The data-logger was designed to be small enough to be instrumented in test models, with an overall size of 100 mm x 100 mm x 70 mm, including batteries. Since the logger was designed to measure force and pressure, it includes six piezoelectric amplifiers and four piezoresistive amplifiers, as well as high-speed analogue-digital converters, which digitize the measured data with 16-bit resolution. The logger’s sampling rate and sample size are 500 kHz and 400 ms, respectively. For the autonomous operation, the logger waits for a trigger signal (accelerometer output) and then starts to take measurements with arbitrary adjustable trigger threshold level and pre-trigger delay time. Measured data is stored to static memory for transfer to a PC via a USB interface after a wind tunnel test. To demonstrate the entire measurement system, wind tunnel experiments were conducted in HIEST. In the present wind tunnel test campaign, records of pressure, axial force, nominal force and pitching moment were obtained under conditions of H0 = 4 MJ/kg, P0 = 14 MPa. This demonstrated that the system worked correctly in the short test duration and harsh conditions typical of HIEST. Use of this data-logger allows the elimination of a large-diameter sting, ending concerns about the sting’s interference with the base flow of the model, which could cause serious errors in measurement in wind tunnel tests.

Effect of Combustor Shape on Scramjet Characteristics at Hypervelocity Condition over Mach 10 Flight

A new scramjet engine model was developed to investigate performance characteristics of scramjet engines at hypervelocity condition over Mach 10 flight. In the design of the new engine, the pressure at the combustor entrance was set to be twice higher than that of the previous engine and the combustor length was reduced while remaining the gas temperature at the combustor entrance adequately low. The combustion test was conducted using a high enthalpy shock tunnel HIEST. The results showed that the combustor performance at the stagnation enthalpy condition of 7MJ/kg or higher was remarkably improved with the present engine. The results positively supported that the high pressure at the combustor entrance and the reduced combustor length was advantageous to achieve high combustor performance at the hypervelocity condition. An attempt was made to apply a combustor with a diverging section to the present engine with an aim of further reduction of the net heat release loss by gradually expanding the combusting gas flow. However, no significant improvement in the combustor performance was observed in the present application.

An experimental study on aerodynamic characteristics of standard model HB-2 in high enthalpy shock tunnel HIEST

Force measurement of a standard model HB-2 was performed in high enthalpy shock tunnel HIEST to study its aerodynamic characteristics. The force measurement results were compared with that obtained in conventional 1.27m hypersonic wind tunnel HWT1. The comparison showed that HIEST results agreed well with that of HWT1 in case of low enthalpy condition. The real gas effect on aerodynamic characteristics was also studied in case of high enthalpy condition.

Numerical Analysis of Experimental Results on the Surface Catalycity in HIEST

Experimental results of a catalytic model in a free-piston driven shock tunnel (HIEST) was analyzed using a non-equilibrium CFD code. Results showed that almost all oxygen molecules were dissociated behind the bow shock and that about 80% of the dissociated oxygen reached the model surface in a typical HIEST flow condition, in the case of noncatalytic wall assumption. In addition to the condition to use air as the test gas, conditions with various oxygen fractions were also tested and compared with numerical analysis. The influence of oxygen fraction in the test gas to the shock-standoff distance agrees well between experiment and numerical analysis. However, the heat flux to the model surface was different between experiment and numerical analysis, especially in the conditions with low oxygen fractions.