Force Measurement Techniques for Hypersonic Flows in Shock Tunnels

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

The experimental activities at Shock Wave Research Center of Tohoku University using a ballistic range are reviewed. First, the shock stand-off distance for a sphere and the shock shape over a sharp cone are measured in the hypersonic speed regime. These measurements are made for 30 mm diameter models, flight speeds between 2.5 and 4.0 km/sec and ambient pressure 4.2 and 76 mmHg. The flow field is visualized by the schlieren method using a short-pulse Nd-YAG laser as the light source. Shock stand-off distance and shock shape are shown to vary with the ambient pressure, demonstrating the effect of the chemical nonequilibrium. Secondly, hypervelocity impact tests are conducted for simulation of space debris impacts against a bumper shield. The bumpers are made of metal and nonmetal materials. The projectile made of high density polyethylene has a cylindrical shape of 14 mm diameter and 3.7 g in weight. The impact velocity is about 4 km/sec. The debris cloud is visualized by shadowgraph. The nonmetal bumper shields used in this study are more effective than the metal bumper in regard to both the bumper mass and defensive capacity. Introduction The ballistic range is one of the readily available impulse facilities for the fundamental investigation of the hypersonic flow phenomena. The so-called two stage light gas gun can accelerate a model to hypersonic speeds. The facility is used mainly for the study of hypersonic aerothermodynamics and for the hypervelocity impact tests. In Shock Wave Research Center of the Institute of Fluid Science of Tohoku University, experiments are in progress in a ballistic range with a 14 mmand 30 mm-bore two stage light gas gun. First, the shock stand-off distance for sphere and the shock shape over sharp cone are measured in the thermochemical nonequilibrium regime. In the past, a large number of experimental investigations on high temperature real-gas effects have been carried out in the shock tunnel. However, there exist uncertainties in the characteristics of the flow produced in a shock * Graduate student f Professor, Institute of Fluid Science, Tohoku University; Member AIAA. Copyright ©1998 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. tunnel in real-gas regime, and therefore it is necessary to validate the facility.Theoretically, the real-gas effects are being studied using computational fluid dynamics (CFD). However, the CFD calculation of nonequilibrium flow encounters many difficulties because of the complex processes of chemical reactions in the nonequilibrium state, particularly in the case of air. Therefore reliable experimental data are needed for validating shock tunnel results and CFD. One of the most appropriate reference quantity is the shock stand-off distance for a sphere and shock shapes for a sharp cone, because the heat absorption by the real-gas phenomena behind shock wave sensitively affects the shock stand-off distance and shock shapes.'' Ballistic ranges are the only impulse facility capable of producing a reliable data, because, in a ballistic range, a model flies in a quiescent atmosphere. Measurement of shock stand-off distance for sphere has been made by Lobb in the ballistic range (see Figure I). His experiment misses the range of the binary scaling parameter pR above 2X10 kg/m and the flight velocity below 4 km/sec. The accuracy of Lobb's experiment is unclear because of the relatively long pulse duration of the light source used for flow visualization and small size of the models. Therefore, we conduct an experiment using a Nd-YAG laser of very short pulse duration for the light source and use 30 mm-diameter models. The shock shapes have been also measured in real gas regime. However, these measurements are conducted in high enthalpy shock tunnels and expansion tubes.' As mentioned above, the data

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.

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.

The diffraction of a shock wave over a stationary body is a problem of interest associated with the starting of shock tubes and expansion tubes, which are well suited to studies of hypersonic magnetohydrodynamic flows. However, these facilities are characterized by very short test times. The transient parameters during the establishment of the detached bow shock in such impulsive facilities are important for both data processing and experimental design. In the present study, based on the low magnetic Reynolds number assumption and dipole magnetic field distribution, the influence of magnetic field on the diffraction of an incident shock wave over a sphere was studied numerically. The incident shock Mach number ranges from 11 to 15 under different magnetic field intensities. Time histories of the shock-detached distance and stagnation pressure were first obtained. Moreover, the time needed to establish the steady flows over the sphere was also displaced against the magnetic interaction parameter. The larger the magnetic interaction parameter, the longer the time needed to establish a steady bow shock for experiments, which further challenges the facilities to have sufficient test times for conducting hypersonic magnetohydrodynamic flow experiments.

In the area of hypervelocity aeroshells, conflicts may be created by the parallel requirements of high aero force braking and stability. This thesis presents a theoretical and experimental study of aerodynamic force coefficients in hypervelocity reacting flows, at conditions of limited static stability. Planetary entry vehicles with sphere-cone shaped aeroshells have been observed in the past to suffer static instability or a reduction of static stability at hypersonic speeds and low angles of attack. This, in many cases, contradicts predictions based on Newtonian theory, may be counter-intuitive, and is an important consideration in vehicle design. Regions of instability along a trajectory can be identified with extensive CFD analysis or experiments, but it is time consimiing and expensive to do so for a complete flight path. Methods for rapidly finding conditions which are predisposed to potential instability are of great value, and enable computational and experimental efforts to be used most effectively. The phenomenon has been found to regularly occur in association with a transition in the sonic line, so that identification of sonic line transition forms a useful indicator of possible areas of concern. Sonic line transition may be initiated by thermo-chemical effects, and the change in the postshock ratio of specific heats has been used successfully in the past to flag potential regions of concern. An improved analytical method is presented here which predicts the location of the sonic line in nonequilibrium flows, taking into account both the chemical composition and effective specific heat of the reacting gas. The approach taken is to model the primary factors that determine the location of the sonic line. Where a transition in the sonic line is seen, the condition can then be targeted with CFD and experimentation to accurately assess stability, as sonic line transition alone does not necessarily imply reduced stability. As validation, the theoretical model was used for a study of the Mars Pathfinder reenti-y path, and correctiy picked up the sonic line transitions known to occur for that trajectory. In addition, a potential instability on the Beagle2 ti-ajectory was also identified. The analytical model was then used to predict a likely transition within conditions typical of the present capabilities of the X3 superorbital expansion tube in a hypersonic flow of Carbon Dioxide. A one piece, 3 component stress wave force balance with a short sting and filleted stress-bars was designed. The ability of the balance to measure small moments in the presence of large drag forces was modelled using Finite Element Analysis, and demonstrated through appropriate calibration. The stress wave force balance was constructed and calibrated, and the sensitivity of the balance to load distribution at appropriate levels and load distributions was quantified. The balance was then used to measure forces on a scale model of the Pathfinder in X3, the largest expansion tube in the world at conditions shown by the analysis to be ones where potential static instabilities may exist for this geometry. The experiments were the first to be conducted in X3. The experiments also are the first measurement of simultaneous axial and normal forces as well as pitching moments in an expansion tube. The measured aerodynamic forces and moments were then used to assess the static stability of the geometry at the conditions tested. This was also the first assessment of static stability in an expansion tube. Results of the experiments indicate that the static stability of the capsule is significantly less than Newtonian theory would predict at the tested conditions. The stress wave balance and the theoretical technique both have applications beyond the sphere-cone geometries on which they have been demonstrated and proven. For instance, any mission where hypersonic stability is important, and where accurate aerodynamic force coefficients are needed may beneficially be studied in impulse facilities using the stress wave force balance. Possible examples are aerocapture vehicles, waveriders and inflatable ballute reentry technology. Some of these applications should be easier to study than the test case used in this work, as the larger values of lift to drag ratio mean none of the measured properties are near zero at the points of interest.

Chapter 11 - Shock Tunnel Studies of High-Enthalpy Ionized Airflows

The shock tunnel ground test is vitally important to the research of the high-enthalpy aerodynamic characteristics of hypersonic vehicles, and the high-accuracy aerodynamic measurement is the key technology. When a force measurement test is conducted in an impulse shock tunnel, the flow field is established instantly after the starting process of shock tunnel, at this time, the great impact loads are acting on the force measurement system. The force measurement system is excited under the action of instantaneous impact, and the inertial vibration signal of the system cannot be rapidly attenuated during the short test time. The output signal of the balance will contain the interference due to the inertial vibration, which leads to a bottleneck in the further improvement of the accuracy of the transient force test. In order to improve the force measurement accuracy in the short-duration shock tunnel, the development of high-accuracy dynamic calibration technology is the key method to improve the performance of balance affected by inertial interference. Therefore, in this paper, recurrent neural network is used to train and intelligently process the balance dynamic calibration data, aiming to eliminate the vibration interference signals in the output dynamic signals. The error analysis of the current method is carried out, and the reliability of the current method is verified. The method is applied to the data processing of force test obtained in shock tunnel, and the effect of inertial vibration on the output signal of the balance is effectively reduced. According to the sample verification analysis of the intelligent model, the relative error of each component load is relatively small, where the case of high-frequency axial force component is about 1%. In the verification of wind tunnel force test data, the good results are also obtained, which are compared with those processed by the convolutional neural network model.

Expansion tubes perform an important role in hypersonic flight research by providing a ground-based testing facility in which scale models can be tested at simulated flight conditions. High enthalpy and high velocity flow conditions are generated via an unsteady expansion process which accelerates test gas to hypersonic conditions. Unfortunately, due to the high velocities and low densities created by this expansion process, test times are shortened and very thick boundary layers develop, reducing the useable core flow. This restricts the size of the models that can be tested, as well as the ability to obtain sufficient or accurate experimental data.Previous studies have shown that a nozzle could be added to an expansion tube to not only expand the core flow diameter but also increase the test time. This included the successful implementation of a nozzle on the University of Queensland's X2 expansion tube [31]. In this thesis, a full-capture, contoured, shock-free nozzle has been designed for the X3 expansion tube. The X3 nozzle contour was scaled from the X2 nozzle contour and was originally 3.0075m long.CFD modelling of the X3 expansion tube with the nozzle attached was completed using a hybrid method coupling a one-dimensional and axisymmetric simulation in order to predict the flow conditions produced by the nozzle. Code package L1d [16] was utilised for the onedimensional analysis and MB_CNS [14], for the axisymmetric analysis. The core flow diameter was increased from 150mm to approximately 250mm and the test time increased by 70% to 1ms; confirming that the addition of a nozzle to the X3 expansion tube will expand the core flow and increase the test time. The nozzle was also examined in relation to its potential to be truncated and maintain its core flow diameter and test time. It was determined that the nozzle could be truncated to a length of 2.5m with no observed loss to the core flow size and test time achieved.The X3 nozzle is under construction. A fibre-wound composite construction technique is being utilised. This construction method was previously used by Jacobs in the fabrication of a nozzle for the T4 reflected shock tunnel [18] at the University of Queensland, although, has not been used for expansion tube nozzles. This technique has proved effective in providing a low-weight, reduced cost alternative to metals.The mandrel for the fibre-winding process was fabricated from high-density polyurethane foam in order to overcome issues of weight and surface instability experienced with other materials. This material had not been previously used for this application and proved very successful in this providing a rigid and stable surface that was suitable for machining as well as winding over.On installation of the nozzle on the facility, the testing of larger, more accurate models is possible. Test times will also be longer than previously observed, enabling the collection of more data, further improving the capacity to effectively model hypersonic flight conditions with this facility. This will benefit future ground based hypersonic flight research to beconducted in the X3 expansion tube facility. Overall, the work completed in this thesis has furthered the development of the X3 expansion tube and provided a foundation for greatly expanding the capabilities of this facility.

The T6 Stalker Tunnel is a multi-mode, high-enthalpy, transient ground test facility. It is the first of its type in the UK. The facility combines the original free-piston driver from the T3 Shock Tunnel with modified barrels from the Oxford Gun Tunnel. Depending on test requirements, it can operate as a shock tube, reflected shock tunnel or expansion tube. Commissioning tests of the free-piston driver are discussed, including the development of four baseline driver conditions using piston masses of either 36 kg or 89 kg. Experimental data are presented for each operating mode, with comparison made to numerical simulations. In general, high-quality test flows are observed. The calculated enthalpy range of the experimental conditions achieved varies from 2.7hbox { MJ kg}^{-1} to 115.0hbox { MJ kg}^{-1}.Graphical abstract

Experimental results for dissociating nitrogen flow over a wedge, obtained in a free-piston shock tunnel, are described. Interferograms of the flow show clearly the curvature of the shock wave and the rise in fringe shift after the shock associated with the dissociation. It is shown that the shock curvature a t the tip of the wedge can be used to calculate the initial dissociation rate and that it is a more sensitive indication of the rate than can be obtained from fringe shift measurements under the prevailing experimental conditions. Because the free-stream dissociation fraction can be adjusted in the shock tunnel, the dependence on atomic nitrogen concentration of the dissociation rate can be determined by the shock curvature method. A detailed calculation of the flow field by an inverse method, starting from the measured shock shape, shows good agreement with experiments.

Shock tunnels create very high temperature and pressure in the nozzle plenum and flight velocities up to Mach 20 can be simulated for aerodynamic testing of chemically reacting flows. However, this application is limited due to milliseconds of its test duration (generally 500 μs–20 ms). For the force test in the conventional hypersonic shock tunnel, because of the instantaneous flowfield and the short test time [1–4], the mechanical vibration of the model-balance-support (MBS) system occurs and cannot be damped during a shock tunnel run. The inertial forces lead to low frequency vibrations of the model and its motion cannot be addressed through digital filtering. This implies restriction on the model’s size and mass as its natural frequencies are inversely proportional the length scale of the model. As to the MBS system, sometimes, the lowest natural frequency of 1 kHz is required for the test time of typically 5 ms in order to get better measurement results [2]. The higher the natural frequencies, the better the justification for the neglected acceleration compensation. However, that is very harsh conditions to design a high-stiffness MBS structure, particularly a drag balance. Therefore, it is very hard to carried out the aerodynamic force test using traditional wind tunnel balances in the shock tunnel, though its test flow state with the high-enthalpy is closer to the real flight condition.

The present paper discusses the base pressure and heat transfer of various blunt bodies at hypersonic speeds. It is shown that the base to stagnation pressure ratio of a blunt body is relatively independent of the Mach and Reynolds numbers in hypersonic flow. The base to stagnation heat transfer, however, shows a dependency on the latter, as it is found to increase with Reynolds numnber. Experimental results of surface heat transfer rate around a cylinder model are presented. These data were obtained in a free piston shock tunnel with total enthalpy of 3.98 MJ/kg. The free-stream unit Reynolds number per metre and Mach number were 1.47 × 106 and 10.9 respectively. The fore-body heat transfer results show good agreement with the theories of Lees and Kemp et al. The base region heat transfer data showed a local minimum at around 130° from the stagnation point followed by a gradual recovery towards the rear stagnation region. The magnitudes of heat transfer ratio in the base region were representative of those in the base region of planetary entry type configurations.

This paper provides a comprehensive discussion of the Mach 5 shock tunnel facility, ST5, located at the Hypersonic Research Center in New Mexico State University. Operation principles of low-enthalpy shock tunnels, as well as the theory and equations used to estimate stagnation properties and free stream conditions of such facilities are reviewed. The material and thickness of the secondary diaphragm are demonstrated to have a significant impact on the steadiness of the stagnation properties. Stagnation properties of the test fluid are determined by analyzing the pressure-time history in the driven tube of the facility. Schlieren flow visualization over a 10deg half-angle straight circular cone and a sphere is employed to estimate the free stream Mach number. Furthermore, post-shock total pressures are measured within the test section using a spherical test model to verify free stream Mach number and test time. The ST5 tunnel has demonstrated the capability to achieve total temperatures ranging from 610-630 K, a free stream Mach number of 5.1, and a test time in the range of 2-2.5 ms. The facility was made available for use in undergraduate courses in Fall 2022.

Measurement of high-bandwidth nanonewton forces in a low-compliance configuration