Shock Tunnel Research Articles

0902 高エンタルピ衝撃風洞での境界層遷移トリップの開発(OS9-1 超音速流れの基礎と応用,OS9 超音速流れの基礎と応用)

Development of boundary layer trip for capsule shaped reentry vehicles under high-enthalpy hypersonic condition was developed in the free-piston shock tunnel HIEST. Wind tunnel test campaign with an 6% scaled Apollo CM model was conducted with stagnation pressure of 30MPa to 60Mpa and with stagnation enthalpy from 3.5MJ/kg to 22MJ/kg. Reynolds number based on model diameter was varied from 0.2 to 1.3 million. To measure heat flux around the model, eighty-four miniature co-axial thermocouples were instrumented on the heat shield surface of the model. Angle-of-attack of the model was fixed to 28degree. A pizza-box configuration boundary layer trip was mounted on the heat shield surface of the model, which trip height k were 0.3mm, 0.6mm and 0.8mm. With k=0.6mm height trip, boundary layer became fully turbulent less than H0=10MJ/kg. However, boundary layer was still laminar over H0=15MJ/kg even with k=0.8mm trip.

Development, Uniformity and Peak Heating of a Hypersonic Laminar Near-Wake Flow

This paper describes visualisation experiments performed using the planar laser-induced fluorescence of nitric oxide. The focus was to investigate the establishment and cross-stream uniformity of the hypersonic separated flow generated in a pulsed free-piston shock tunnel facility. In particular, the laminar Mach 7.5 flow over a step downstream of a wedge was investigated. The ratio of wedge width to total horizontal distance downstream of the leading edge was 0.71 at the furthest downstream measurement point, and model side plates were not used. Flow images were obtained on successive tunnel runs with different delays after flow onset, to identify the nominal flow establishment time. Furthermore, images were obtained at the nominal facility test time using a laser sheet oriented perpendicular to the step, propagating in the spanwise direction at a height equal to half the step height. These images are compared with thermocouple heat flux measurements obtained both with and without the step and show that ...

Estimation of Skin Friction Drag on a Model in Hypersonic Shock Tunnel

Viscous drag on the internal surfaces of a notional scramjet engine model has been estimated through Reynolds analogy, using measured wall heat transfer rates, in a shock tunnel at a hypersonic Mach number of 8. The study has been carried out without fuel injection and at zero degree angle of incidence of the model with the freestream. The heat transfer rate measurements were carried out on the upper and lower internal surfaces of the engine employing fast response E-type thermocouples. Application of Reynolds analogy to the wall heat transfer rates yielded the skin friction coefficient, through which the viscous drag on the surfaces could be determined. The measurements predict the salient features of the flow field of the model and are a novel reference on the data to the researchers working in the area of slender-body, hypersonic aerothermodynamics.

Improvement of a Model Scramjet Engine Design Using Various Types of Ground Tests

The scramjet, one of the most promising candidates for future transport systems, has many advantages such as a simple configuration and a high specific impulse. However, various characteristics of scramjet engines in a hypersonic environment are difficult to predict from theory. In the present study, we investigated the characteristics and the performance of the scramjet engine with various types of ground tests. An engine model (S1) was tested in the free-piston shock tunnel T4. (S1) showed active supersonic combustion and good starting characteristics, but its operating range is rather limited. To improve the performance of the model scramjet engine, we also investigated major components, such as the combustor and the intake by experiments with different model configurations. In the combustor test results, the zigzag cavity flame holder showed the best combustion performance among the three cases. By testing various intake configurations, the most efficient intake configuration was also determined. A revised model (S2), which was redesigned based on the component test results, showed enhanced thrust performance and a wider operating range in the free jet test using the free piston shock tunnel HIEST.

NO PLIF imaging in the CUBRC 48-inch shock tunnel

Nitric oxide planar laser-induced fluorescence (NO PLIF) imaging is demonstrated at a 10-kHz repetition rate in the Calspan University at Buffalo Research Center’s (CUBRC) 48-inch Mach 9 hypervelocity shock tunnel using a pulse burst laser–based high frame rate imaging system. Sequences of up to ten images are obtained internal to a supersonic combustor model, located within the shock tunnel, during a single ~10-millisecond duration run of the ground test facility. Comparison with a CFD simulation shows good overall qualitative agreement in the jet penetration and spreading observed with an average of forty individual PLIF images obtained during several facility runs.

A computational study of the HyShot II combustor performance

Experimental and flight data for hypersonic air-breathing engines are both difficult and extremely expensive to obtain, motivating the use of computational models to enhance the understanding of the complex physics involved. Here, a comprehensive numerical study has been carried out for the HyShot II scramjet combustor. This study makes use of Reynolds Average Navier Stokes (RANS) based models and Large Eddy Simulation (LES) based models with semi-detailed reaction kinetics. In this investigation we focus on the underlying flow-mixing-combustion physics at different operating conditions tested in the High Enthalpy Shock Tunnel Göttingen (HEG). To account for the complex flow in the HEG facility a zonal approach is employed in which RANS is used to simulate the flow in the HEG nozzle and test-section, providing the necessary inflow boundary conditions for the combustor RANS and LES, being the focus of this analysis. Specifically, we here combine results from RANS and LES computations with data from the HEG experiments and the target HyShot II flight-tests at two different flight-altitudes (28 and 33km). The LES model is observed to capture the experimental wall-pressure and heat-flux data very well for both the 33 and 28km altitude cases, whereas the RANS model is only able to predict the wall-pressure and heat flux data for the 28km altitude case. Based on the LES results, the flow at both altitudes is found to be unsteady, but with unsteady transitional flow features dominating the 33km case. Moreover, these results show that the equivalence ratio is of key importance to the resulting flow, mixing and combustion physics, with richer mixtures being prone to transitional flow features. The LES results are also used to describe the flow physics in detail for both altitudes, and the key processes responsible for the transition between the two combustion modes observed.

Force Measurements in Hypervelocity Flows with an Acceleration Compensated Strain Gage Balance

Force and moment measurements are performed on a 6-inch Apollo shaped capsule in the LENS I and 48-Inch reflected shock tunnels at CUBRC in cold hypersonic nitrogen flows and for high enthalpy flows (10 MJ/kg) of air, nitrogen, pure oxygen and a 50%-50% oxygen-argon mixture. Measurements with an acceleration compensated flexured strain gage force balance are compared with non-equilibrium Navier-Stokes computations. For the low enthalpy nitrogen runs, axial and normal force coefficients are within 2% of the numerical predictions whereas the pitching moment is under-predicted by 10%. Excellent agreement within 0.5% for the lift-to-drag ratio is achieved. For the first time, the effect of thermodynamic and chemical non-equilibrium on the force and moment coefficients is measured on a reentry capsule in a reflected shock tunnel and compared with state of the art numerical simulations. Both computations and experiments show an increase in the axial force, normal force and pitching moment coefficients and a decrease in the lift-to-drag ratio.

Measurement of yaw, pitch and side-force on a lifting model in a hypersonic shock tunnel

Yawing moment, pitching moment and side-force on a blunt, triangular, lifting model were measured in a shock tunnel using an accelerometer balance at a hypersonic Mach number of 8. The measurements were carried out at different angles of incidence of the model with the freestream. The balance was instrumented with accelerometers to sense the model acceleration in the desired directions, and was equipped with a soft suspension system such that the model had a free flight during the tests in the shock tunnel. Pressure measurements on the surfaces of the model were carried out using high frequency pressure transducers to validate the output of the force balance. The moment and force coefficients obtained through both the measurement techniques have a good agreement reciprocally. The force and moment coefficients for the test model were analytically estimated using Newtonian theory to compare with the measured coefficients. The comparison between theory and experiments has been found quite encouraging.

Experimental study of shock tunnel flow with a stationary throat plug

A throat plug is a device to prevent fragments produced by the bursting of the primary diaphragm of a shock tunnel from entering the nozzle and damaging the model. An experimental study has been carried out to investigate the flow in the KAIST shock tunnel with two stationary throat plugs at primary shock velocities of 1.19 and 1.28 km/s. These shocks generate tailored conditions when helium and air are used as the driver and driven gases, respectively, at room temperature. Nozzle reservoir pressures and Pitot pressures at the nozzle exit are measured to examine the influence of the stationary throat plug on these properties. The throat plug that has an areal blockage of 6.2 % is found to have negligible effects on shock tunnel flow. However, the throat plug of 19.4 % blockage shows an appreciable influence by generating a pressure bump. Although it retards the establishment time of the nozzle flow, after the transient period, it does not cause deviation of the pressure from that for the clean nozzle configuration. It is found that the reduction of the cross-sectional area, which results from the presence of the throat plug, causes the pressure bump. By increasing the diameter of the driven tube in the region where the throat plug is located appropriately, the magnitude of the pressure bump is reduced significantly.

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.

S191014 実在気体気流でのHRVカプセル空力加熱試験

Aerodynamic heating test has performed in the free-piston high-enthalpy shock tunnel JAXA-HIEST. A HRV 6% scale model has applied in the test, which model dimension was decided as a reference configuration of HRV. The model has eighty-eight fast-response coaxial thermocouples on the windward surface to measure heat flux distribution. Thirty-four thermocouples and eight Piezo-resistive pressure transducers were also instrumented on the aft of the model and were used to determine flow establishment around the model through test period. Stagnation enthalpy and stagnation pressure were varied from H_0=3MJ/kg to 25MJ/kg, and from P_0=14MPa to 52MPa, respectively. The unit Reynolds number was changed from 0.9 million/m to 4.5 million/m. Aeroheating characteristics were observed at model angle of attack 0 degree and 20 degree with fully laminar or with transitional boundary layer. Measured heat flux was normalized by the products of the Stanton number and the square-root of the Reynolds number for comparison with other wind tunnel test results. In the present test campaign abnormally high heat-flux phenomenon was observed, which phenomenon appeared under stagnation enthalpy H_0 > 13MJ/kg and stagnation pressure P_0 > 45MPa

Preliminary Experimental Investigation on MHD Power Generation Using Seeded Supersonic Argon Flow as Working Fluid

This paper presents a preliminary experimental investigation on magnetohydrodynamic (MHD) power generation using seeded supersonic argon flow as working fluid. Helium and argon are used as driver and driven gas respectively in a shock tunnel. Equilibrium contact surface operating mode is used to obtain high temperature gas, and the conductivity is obtained by adding seed K2CO3 powder into the driven section. Under the conditions of nozzle inlet total pressure being 0.32 MPa, total temperature 6 504 K, magnetic field density about 0.5 T and nozzle outlet velocity 1 959 m/s, induction voltage and short-circuit current of the segmentation MHD power generation channel are measured, and the experimental results agree with theoretical calculations; the average conductivity is about 20 S/m calculated from characteristics of voltage and current. When load factor is 0.5, the maximum power density of the MHD power generation channel reaches 4.797 1 MW/m3, and the maximum enthalpy extraction rate is 0.34%. Finally, the principle and method of indirect testing for gas state parameters are derived and analyzed.

Investigations on supersonic combustion in a combustion-driven shock tunnel

This article reports experimental results on supersonic combustion in a new facility. The facility is a combustion-driven shock tunnel, which is cheaper to build than the facilities in which such experiments are carried out conventionally. The observation region is a zone between two parallel flat plates with a 33° wedge attached to the upstream end of the bottom plate. Gaseous hydrogen is injected at an angle of 45° into an oncoming supersonic flow of Mach 2 (approximate) from a port on the bottom plate. The resulting flow field is visualized by a high speed camera in a dark background. Three different test gases, namely nitrogen, air, and oxygen-rich air are used, and the results are compared. A distinct luminosity due to combustion for oxygen-containing test gases is observed. Heat-transfer rates on a probe placed at the downstream end of the observation region and midway between the parallel plates are measured and compared for the three cases. Wall static pressure at 28 mm downstream of the injection port on the bottom plate is also presented.

OH-PLIF visualisation of radical farming supersonic combustion flows

Experimental investigations employing Planar Laser-induced fluorescence visualisation of the qualitative distribution of the OH radical (OH-PLIF), coupled with surface pressure measurements, have been made of flow in a generic, nominally two-dimensional inlet-injection radical farming supersonic combustion scramjet engine model. The test flows were provided by a hypersonic shock tunnel, and covered total enthalpies corresponding to the flight Mach number range 8.7–11.8 and approximately 150 kPa dynamic pressure. The surface pressure measurements displayed radical farming behaviour, that is a series of adjacent high and low pressure regions corresponding to successive shock/expansion structures, with no significant combustion-induced pressure rise until the second structure. OH-PLIF imaging between the first two structures provides the first direct experimental evidence of significant OH radical concentrations upstream of the ignition point in this mode of scramjet operation and shows that combustion reactions were occurring in highly localised regions in a complex turbulent and poorly micromixed fuel/air mixing layer confined to the fuel injection side of the combustor.

On tracking the motion of rigid bodies through edge detection and least-squares fitting

A class of techniques is investigated for determining at least three components (two translational and one rotational) of the motion of a rigid body from silhouette images, with particular emphasis on motion within a fluid. The investigated techniques all employ edge detection followed by some form of least-squares fitting to the detected points in determining the movement of the body. Four techniques are discussed and, through both an artificial image analysis and calibrated sphere measurements, are shown to be capable of measuring displacements down to a few thousandths of a pixel under low image-noise conditions (\(\lesssim2\%\)). Measurements of two configurations in a high-enthalpy shock tunnel demonstrate the capabilities of the techniques under experimental conditions. In particular, a technique referred to as edge-tracking is introduced, which can be employed in situations where the model profile is unknown and/or only some fraction of it is visible. This latter quality is especially useful for measurements in high-enthalpy facilities, where test-gas luminosity can obscure a significant extent of the model outline. A further advantage of this technique is that, even for complex geometries, the fitting procedure can typically be reduced to solving a sequence of linear least-squares problems, rather than a nonlinear one, with a corresponding benefit in computational efficiency.

Cowl and Cavity Effects on Mixing and Combustion in Scramjet Engines

To investigate the supersonic combustion patterns in scramjet engines, a model scramjet engine was tested in the T4 free-piston shock tunnel. The test model had a rectangular intake, which compressed the freestream flow through a series of four shock waves upstream of the combustor entrance. A cavity flame holder was installed in the supersonic combustor to improve ignition. The freestream test condition was fixed at Mach 7.6, at an altitude of 31 km. This experimental study investigated the effects of varying fuel equivalence ratios, the influence of the cavity flame holder, and the effects of cowl shape. As a result, supersonic combustion was observed at equivalence ratios between 0.11 and 0.18. Measurements indicated that the engine thermally choked at a fuel equivalence ratio of 0.40. Furthermore, the cavity flame holder and the W-shaped cowl showed improved pressure distribution due to greater reaction intensity. With the aid of numerical analysis, the cavity and the W-shaped cowl are shown to be effective in fuel–air mixing.

EXPERIMENTAL RESULTS OF A MACH 10 CONICAL-FLOW DERIVED WAVERIDER TO 14-X HYPERSONIC AEROSPACE VEHICLE. doi: 10.5028/jatm.2011.03027510

This paper presents a research in the development of the 14-X hypersonic airspace vehicle at Institute for Advanced Studies (IEAv) from Department of Science and Aerospace Technology (DCTA) of the Brazilian Air Force (FAB). The 14-X project objective is to develop a higher effi cient satellite launch alternative, using a Supersonic Combustion Ramjet (SCRAMJET) engine and waverider aerodynamics. For this development, the waverider technology is under investigation in Prof. Henry T. Nagamatsu Aerothermodynamics and Hypersonics Laboratory (LHTN), in IEAv/DCTA. The investigation has been conducted through ground test campaigns in Hypersonic Shock Tunnel T3. The 14-X Waverider Vehicle characteristic was verifi ed in shock tunnel T3 where surface static pressures and pitot pressure for Mach number 10 were measured and, using Schlieren photographs Diagnostic Method, it was possible to identify a leading-edge attached shock wave in 14-X lower surface.

Hypervelocity Experiments on Oxygen Enrichment in a Hydrogen-Fueled Scramjet

Extension of scramjet operation into the hypervelocity regime (u > 3 km/s, M > 10) is critical for their use in access-to-space systems. As flight speeds increase, the net thrust of a scramjet engine is limited by increased fuel/air mixing length requirements, as well as the decreased ratio of energy available from stoichiometric fuel/air combustion relative to the specific kinetic energy of the captured air. One possible method for addressing these issues, known as oxygen enrichment, involves the premixing of oxygen with fuel before injection. This method provides a head start for mixing and also allows the possibility of fuel combustion at greater than stoichiometric proportions. This paper describes some exploratory shock tunnel experiments involving oxygen enrichment of gaseous hydrogen fuel in a scramjet at hypervelocity conditions. These experiments were conducted in a simple rectangular duct with a central strut injector and varying levels of enrichment. The measured pressure distributions show a positive effect of enrichment and indicate how this technique could be best used to extend the performance of hypervelocity scramjets to higher flight Mach number. Copyright

An accelerometer balance for the measurement of roll, lift and drag on a lifting model in a shock tunnel

A force balance to measure roll, lift and drag on a lifting aerodynamic body in an ultrashort-duration hypersonic test facility, such as a shock tunnel, has been developed and tested on a flapped, blunt-nosed, triangular lifting body at a freestream Mach number of 8. The flow total enthalpy and the freestream unit Reynolds number were 0.83 MJ kg−1 and 0.98 million, respectively. The balance structure has a soft suspension that allows the model to have a free flight during the short-duration aerodynamic test. The balance was mounted inside the hollow model and was equipped with accelerometers to sense the aerodynamic moment and forces on the model. The measurements were carried out at different angles of incidence of the model and the acquired signals of the accelerometers were reduced to the aerodynamic moment and the force coefficients based on the theories of applied mechanics and aerodynamics. Also, the moment and force coefficients were theoretically calculated based on the Newtonian theory, which is an accepted analytical approach for hypersonic bodies. Good agreement has been observed between the experimental and the analytical results. The method of measurement of roll and lift, and the data on the rolling moment of a lifting body presented in this note are novel.

Application of particle image velocimetry and the background-oriented schlieren technique in the high-enthalpy shock tunnel Göttingen

The applicability of the particle image velocimetry (PIV) and the background-oriented schlieren (BOS) techniques in the high-enthalpy shock tunnel Gottingen of the German Aerospace Center, DLR is demonstrated. As a part of this feasibility study two different experiments are performed. The velocity field past a wedge in a Mach 6 flow at a total specific enthalpy of 1.5 MJ/kg is determined by means of PIV and the results are compared to numerical predictions. The BOS technique is applied to investigate the density field in the shock layer of a sphere at 12 and 22 MJ/kg total specific enthalpies. Using a ray tracer method, the BOS results are compared to the data obtained by corresponding numerical computations.