Overview of ballistic range program at Tohoku University

An experimental demonstration was carried out in a ballistic range at high Mach numbers with the low specific heat ratio gas hydrofluorocarbon HFC-134a to observe the unstable bow-shock wave generated in front of supersonic blunt objects. The shadowgraph images obtained from the experiments showed instability characteristics, in which the disturbances grow and flow downstream and the wake flow appears wavy because of the shock oscillation. Moreover, the influence of the body shape and specific heat ratio on the instability was investigated for various experimental conditions. Furthermore, the observed features, such as wave structure and disturbance amplitude, were captured by numerical simulations, and it was demonstrated that computational fluid dynamics could effectively simulate the physical instability. In addition, it was deduced that the shock instability is induced by sound emissions from the edge of the object. This inference supports the dependence of the instability on the specific heat ratio and Mach number because the shock stand-off distance is affected by these parameters and limits the sound wave propagation.

Hypervelocity impact tests with bumper walled structure are conducted in a two stage light gas gun. The bumpers are made either of aluminum alloy, carbon/polyimide, flexible insulation + Al, 2.5D carbon/carbon + Al, fiber metal laminates, or ceramic. The impact velocity is about 4 km/sec. The debris cloud is visualized by the shadowgraph method or by using a multi-frame image converter camera. The damages to the main wall are compared among the bumper materials. The efficiency of the nonmetal bumper shielding is found to be superior to that of the metal bumper in both specific density and effectiveness protection.

Analysis of an internally mountable accelerometer balance system for use with non-isotropic models in shock tunnels

Many previous studies have addressed the problem of theoretically approximating the shock standoff distance; however, limitations to these methods fail to produce excellent results across the entire range of Mach numbers. This paper proposes an alternative approach for approximating the shock standoff distance for supersonic flows around a circular cylinder. It follows the philosophy that the “modified Newtonian impact theory” can be used to calculate the size of the sonic zone bounded between the bow shock and the fore part of the body and that the variation of the said zone is related to the standoff distance as a function of the upstream Mach number. Consequently, a reduction rate parameter for the after-shock subsonic region and a reduction rate parameter for the shock standoff distance are introduced to formulate such a relation, yielding a new expression for the shock standoff distance given in Equation (32). It is directly determined by the upstream Mach number and the location of the sonic point at the body surface. The shock standoff distance found by this relation is compared with the numerical solutions obtained by solving the two-dimensional inviscid Euler equations, and with previous experimental results for Mach numbers from 1.35 to 6, and excellent and consistent agreement is achieved across this range of Mach numbers.

Heat loads on spacecraft traveling at hypersonic speed are of major interest for their designers. Several tests using temperature-sensitive paints (TSP) have been carried out in long duration shock tunnels to determine these heat loads; generally paint layers were thin, so that certain assumptions could be invoked to enable a good estimate of the thermal parameter ρck (a material property) to be obtained—the value of this parameter is needed to determine heat loads from the TSP. Very few measurements have been carried out in impulse facilities [viz. shock tunnels such as the High Enthalpy Shock Tunnel Gottingen (HEG)], where test times are much shorter. Presented here are TSP temperature measurements and subsequently derived heat loads on a ramp model placed in a hypersonic flow in HEG (specific enthalpy h 0 = ~3.3 MJ kg−1, Mach number M = 7.4, temperature T ∞ = 277 K, density ρ ∞ = 11 g m−3). A number of fluorescence intensity images were acquired, from which, with the help of calibration data, temperature field data on the model surface were determined. From these the heat load into the surface was calculated, using an assumption of a 1D, semi-infinite heat transfer model. ρck for the paint was determined using an insitu calibration with a Medtherm coaxial thermocouple mounted on the model; Medtherm ρck is known. Finally presented are sources of various measurement uncertainties, arising from: (1) estimation of ρck; (2) intensity measurement in the chosen interrogation area; (3) paint time response.

The code calibration program for the Langley Aerothermodynamic Upwind Relaxation Algorithm to be used as support for the Aeroassist Flight Experiment (AFE) is discussed. Comparisons between experimental data and numerical simulations are made which focus on perfect-gas tests involving a scale model of the AFE. Aspects of the thermochemical nonequilibrium model are called into question by the results of ground tests performed in a ballistic range and in a shock tunnel.

Aerodynamic force measurement in high-enthalpy flow is very important for the design and optimization of hypersonic vehicles. Currently, impulse facilities are used for generating high-temperature and high-pressure driving gas to simulate the high-enthalpy flow with hypersonic flight-conditions, such as a shock tunnel. However, when force tests are conducted in an impulse facility, the inertial force has a large influence on the measuring results, which creates low-frequency vibrations of the test model and its motion cannot be addressed through digital filtering. In the case of a few milliseconds of test time, the structural design of the six-component balance is greatly challenged. Therefore, dynamic calibration becomes very important for improving the precision and accuracy of force measurement during short-duration. A new method, deep-learning-based single-vector dynamic self-calibration of the force measurement system, and intelligent force measurement system are proposed for obtaining high-accurate aerodynamic force in impulse facilities. One of the main features of this dynamic calibration method is the calibration of the overall force measurement system, not just the balance. Applying this method, the calibrated force measurement system is the wind tunnel test object, which ensures the consistency of calibration and application. In the evaluation, the test verification has achieved relatively ideal results, the large-scale low-frequency vibration interference has been basically eliminated, and the accuracy and reliability of the force measurement in impulse facility have been greatly improved.

On 5 May 1996 the Interball‐1 and IMP 8 spacecraft crossed the bow shock boundary. The upstream conditions were special in two factors: (1) the interplanetary magnetic field was anti‐parallel to the solar wind flow within 15° and (2) the conditions were stable for a prolonged period (∼9 hours). At the nose of the magnetosphere, the Interball‐1 data revealed that the magnetopause was farther outward by ∼2 RE than model predictions and the subsolar magnetosheath was unusually thin, at most 10% of the magnetopause standoff distance. Both results stand in contrast to predictions of existing magnetopause/bow shock models. Assuming a hyperboloidal (paraboloidal) shock wave, the calculated shock's standoff distance was 13.7 (13.6) RE, and the focus was located on the x axis at 4.5 (4.2) RE. On the basis of the IMP 8 observation, the bow shock flares significantly less than MHD simulations predict for a field‐aligned bow shock at the magnetospheric flanks. This study discusses differences between the observations and existing MHD bow shock simulations for field‐aligned upstream flow. Furthermore, it is suggested that the flow‐aligned IMF orientation causes a significant change of the magnetopause shape into a bullet‐like obstacle.

Recent advances in hypersonic research activities have led to the development in ground based test facilities for simulating near-realistic flight conditions in the laboratory. Basic performance data such as drag, lift, thrust etc., are usually obtained from ground-based experimentation for physical understanding of the flow fields and subsequent prototype developments. Hypersonic flows are produced for a very short time in impulse facilities such as shock tunnels, free-piston shock tunnels and expansion tubes. The test times are in the order of few milliseconds in the shock tunnel and for the expansion tubes, it is still less (~50 μs). The force measurements in these facilities are very challenging due the need of fast response sensors and instrumentations. In this review paper, two force measurement techniques are discussed that are best suited for short duration facilities namely; accelerometer balance and stress-wave force balance system.

A stress wave internal force balance for the High Enthalpy Shock Tunnel Gottingen (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° to 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.

Shock tunnels are important ground test facilities that can generate high-enthalpy flow. Flight velocity at a high Mach number can be simulated for aerodynamic testing of chemically reacting flows. However, the application of these tunnels is limited due to the only milliseconds-long test duration, especially for aerodynamic force measurement using traditional strain gauge balances. This study presents an impulse force-measurement system, which was used for a large-scale test model to measure its drag in a high-enthalpy shock tunnel with an approximately 3–7-ms test time. Force tests were conducted for a cone in the JF-10 high-enthalpy shock tunnel in the Institute of Mechanics, Chinese Academy of Sciences. An integrated design of the impulse force-measurement system was proposed for load measurement over a short duration, for which a recommended design criterion is that the measurement period be a minimum of twice the period corresponding to the lowest natural frequency of the measurement system. The current measurement technique breaks the limitations of the application of the conventional strain gauge balance. As an integrated measuring system, the impulse force-measurement system expands the structural design concept of strain gauge balances. The impulse force-measurement system performed well in the present tests. The test results show differences from the numerical simulations and some data obtained in a conventional wind tunnel. A preliminary analysis was performed on the real gas effects on the aerodynamic force.

Double cone is a commonly used axial symmetric configuration in numerical and experimental studies to simulate complex flow with shock/shock and shock/boundary layer interactions [1–4]. For hypersonic vehicle thermal protection system design, interactions between the bow shock with the shock waves generated by the wing or control surface and between shock wave and boundary layer must be considered of primary concern because it may cause severe local heating load. Double cones with different cone corners become a good choice for aerothermodynamic researchers to investigate the complex heating distribution caused by different kinds of shock interactions. It is critical to accurately determine the spatially continuous heat flux distributions on double cone model surface for understanding of complex flow mechanisms and for comparison with computational fluid dynamics (CFD) predictions, but it is difficult to get the detailed aerodynamic heating data with traditional point measurement techniques, which use separately located thin film gauges or coaxial thermocouples to get the heat flux value on the point. Temperature-sensitive paint (TSP) technique, a global heating measurement technique, makes acquiring such data possible. In this work, a new developed fast-responding TSP system [5] is used to measure the global heat flux distributions on two double cone models in shock tunnel, and the results are compared with the data from thin film gauges and CFD predictions. The main objective of this experiment is to develop experience in the application of TSP in the shock tunnel testing environment and to evaluate the effectiveness of the different measurement and computation techniques in the complex flow field.

A two-dimensional hypersonic scramjet inlet has been investigated in a combined experimental and analytical program aimed at addressing the fundamental issues related to the design of scramjet inlets. The experimental portion of the program was conducted in the Calspan 48-in. shock tunnel at Mach 10 and 13. The computational analysis was conducted using a two-dimensional parabolized Navier-Stokes computational fluid dynamics (CFD) code. This article addresses the issues concerned with the flow over the external forebody of the inlet which consists of a blunted wedge followed by an isentropic compression. The pressure and heat transfer distributions over the forebody are investigated for ranges of Reynolds number, Mach number, wall-to-freestream temperature ratio, and nose bluntness. Comparison of the test results and CFD predictions show that good agreement for the heat transfer distributions is achieved. However, the predicted pressure distribution on the forward blunted wedge was consistently underpredicted by 18-33% relative to the experimental measurements. Several phenomena were investigated in an attempt to explain the discrepancy between predicted and measured pressure distributions, including classical viscous leading-edge interactions, blunt leading-edge interactions, slip flow effects, flow condensation, flow angularity, and facility Mach number uncertainty. Although the discrepancy in the forebody pressure ratios could be caused by a combination of the factors listed above, a deficiency in the modeling of the viscous interaction region by the CFD codes and facility flow angularity are shown to be the strongest contributors.

Effects of controlled heat addition into the high temperature, chemically reacting shock layer of a large angle (60°) blunt cone with a spherical nose have been experimentally investigated. The exothermic oxidation of ablated chromium from the surface of the cone at hypersonic Mach numbers triggers the heat release process. A conical skirt with a base diameter of 70 mm and a semi-apex angle of 30°, culminating into a nose of radius 30 mm, has been used as the test model. An in-house hypersonic free-piston driven shock tunnel facility, HST3, was used for the experiments at a stagnation enthalpy of 6.31 MJ/kg and a freestream Mach number of 9.84. The temperature distribution in the shock layer was experimentally measured by the two-color ratio pyrometry technique, using a Digital Single Lens Reflex (DSLR) camera as a pyrometer. The temperature field was corrected for discrete radiative line emissions obtained through emission spectroscopy. Surface heat flux measurements were taken using carefully calibrated thin film platinum heat transfer gauges mounted on an insulating Macor substrate. Shock stand-off distances were measured through Schlieren imaging of the flow using a high-speed camera at 20 000 frames per second, and also by a new intensity-scan based method using the processed color image from the DSLR camera. Calculations showed a 173 K rise in the temperature of the gas layer in the stagnation region due to chromium oxidation. The net surface heat flux on the blunt cone was also found to increase by about 31 W/cm2. The shock stand-off distance, as ascertained from Schlieren images, increased from about 3.82 mm (±1.4%) to 4.45 mm (±1.5%), a 17% rise. Analytical calculations, taking chromium oxidation reaction kinetics into consideration, to relate the total exothermic heat release to its distribution into various processes demonstrated that chromium oxidation releases about 78 W/cm2 energy in the stagnation region of the shock layer. 1.9% of this energy increases the temperature of the gas layer, 40% is convected back into the airframe, 0.4% is lost from the rear of the Macor substrate by conduction, and about 8% is radiated into the model. The remaining 50% of the heat was used up in pushing the shock layer away from the body by raising its density, thereby increasing shock stand-off distance.

Meaurement of mass. spray angle and velocity distribution of fragment cloud