Hypersonic Test Research Articles

Accelerated pump tube concept for hypersonic ground testing

This paper examines a new hypersonic wind tunnel concept which is referred to here as an ‘accelerated pump tube’ (APT). The facility can be considered a hybrid between a gun tunnel and an expansion tube. Like a gun tunnel, the test gas is isentropically compressed with a piston to raise its pressure and temperature; like an expansion tube, the compressed test gas is subject to an unsteady expansion to further increase its enthalpy and stagnation pressure. State-to-state analytical calculations incorporating high temperature equilibrium chemistry are presented which demonstrate that the APT is theoretically able to simulate a wide hypersonic terrestrial flight envelope from Mach 5–14, with significant capacity for pressure-length scaling. Potential practical benefits of this concept include relatively simple internal flow processes, removal of the need for a film secondary diaphragm and helium as a driver gas, and a reduction in experimental uncertainty through better defined internal boundary conditions and relatively low mechanical complexity. To assess the feasibility of the APT concept in a practical context, the paper considers how an APT based around a 25 m long, 600 mm diameter pump tube should be configured for a Mach 10 test flow. One-dimensional Lagrangian CFD analysis is used to compute piston dynamics for three different pump tube compression ratios and the associated test flows are characterised. The CFD results suggest that the concept should be viable in practice, and also validate the Mach 10 state-to-state calculations.

Representation of Thermomechanical Damage in Fiber-Reinforced Ceramic Composites at High Thermal Gradients

Ceramic matrix composites (CMCs) are the core strategic materials for the thermal structure of new-generation hypersonic aircraft. The unclear understanding of the cross-scale thermomechanical behavior and failure mechanism of materials and structures in high-temperature transient environments is a key scientific issue that restricts the reliable design and safe operation of hypersonic aircraft. In this study, a high-temperature test module in a supersonic wind tunnel with integrated various test systems was independently designed to systematically investigate the thermal mechanical damage evolution of CMCs used in astronautics applications under high-temperature transient conditions. Furthermore, a numerical method for fluid-heat-solid coupling based on an iterative solution strategy was developed and compared with an algorithm based on a wall modification strategy in terms of computational efficiency and accuracy. Besides, a comparative analysis was conducted with the temperature field distribution of the sample obtained in the hypersonic wind tunnel integration test to verify the correctness of the developed thermal-fluid-solid coupling calculation method. Finally, through thermoelastic coupling computational models, the high-thermal-gradient-induced thermomechanical damage mechanism of the CMCs was revealed, providing support for further optimizing the material system and thermal structure design of the CMCs.

Development of force measurement technique for preheated model in hypersonic combined test facility

Development of force measurement technique for preheated model in hypersonic combined test facility

Development of combined hypersonic test facility for aerothermodynamic testing.

In this study, a combined test facility was developed using a combination of an arc-jet tunnel and a shock tunnel for aerothermodynamic testing. The performance validation of individual parts was performed, and results were obtained from the combined test. A small-scale Huels-type arc-jet tunnel was used to preheat the test model by aerodynamic heating before conducting the experiments in the shock tunnel to duplicate the hot surfaces of flight objects encountered during hypersonic flight. The high-enthalpy flow in the arc-jet tunnel provided a heat flux of 1.99±0.03 MW/m2 for a flat-faced model of 10 mm diameters, and the flow condition of the shock tunnel used in this study simulated a Mach 5 flight at a pressure altitude of about 24 km. The two combined experiments employing different shape and material models were carried out to examine the effect of aerothermodynamic phenomena. In the first experiment, the effect of ablation-induced shape change on the fluid-structure was investigated using a cone model manufactured of AL6061 material. The effect of surface roughness on the fluid-structure was examined in the second experiment, which used a hemisphere model constructed of STS303 material. Although substantial findings could not be validated due to the limits of qualitative evaluations utilizing visualization methods, however preheating-related changes in surface roughness were found. As a follow-up study, a force measuring experiment based on the test procedures is being carried out at this facility utilizing a preheated model with an accelerometer. The performance and experimental results obtained using this integrated setup are discussed in detail, highlighting the potential of this combined hypersonic test facility.

High-Sensitivity Atomic Layer Thermopile Heat- Flux Sensor and Its Application in Hypersonic Low-Density Wind Tunnel Tests

High-Sensitivity Atomic Layer Thermopile Heat- Flux Sensor and Its Application in Hypersonic Low-Density Wind Tunnel Tests

Investigation of visible to near-infrared high-speed cameras for thermography in hypersonic testing

Investigation of visible to near-infrared high-speed cameras for thermography in hypersonic testing

Study on Ground Experimental Method of Stagnation Point Large Heat Flux of Typical Sharp Wedge Leading Edge Structure

In this paper, aimed at the problem of large temperature gradient thermal testing with the typical sharp wedge leading edge structure of a hypersonic vehicle, a subsonic high-temperature combustion gas heating (SHCH) test device is used to conduct a series of experiments on the heat flux simulation ability of subsonic high-temperature combustion gas in the stagnation point region. Firstly, for a hypersonic vehicle with a flying height of 24 km and Mach number range of 4~6.5, the stagnation point heat flux in the head area is obtained by numerical calculation of a typical leading edge structure, which is used as the experimental target of the thermal structure test. Secondly, an experimental specimen with a Gardon heat flux meter is designed with the same shape and size as the specimen in the numerical simulations to prepare for the subsequent SHCH test. Thirdly, a method to determine the combustion gas temperature based on a Kriging surrogate model is proposed. CFD numerical simulation is conducted using the SHCH test model, and the numerical calculation results are used as the training dataset. The Kriging surrogate model is used to establish an approximate fitting relationship between the stagnation point heat flux and experimental parameters under SHCH conditions. The corresponding combustion gas temperature values are found, respectively, with the hypersonic aerodynamic heat flux at Mach 5.0~5.4 as the target value. Finally, stagnation point heat flux testing of low-speed and high-temperature combustion gas is performed at different combustion gas temperatures. The experimental and target values obtained from hypersonic aerodynamic thermal simulations are compared and analyzed to verify the heating capacity of SHCH and the feasibility of hypersonic aerothermal simulation testing.

Nonlinear constitutive calculation method of rarefied flow based on deep convolution neural networks

In the field of rarefied gas dynamics, the presence of non-equilibrium flow characteristics poses significant challenges for achieving efficient and accurate numerical simulation methods. These challenges arise from the complex coexistence of these phenomena at multiple scales. The recent advent of intelligent fluid mechanics has introduced the data-driven nonlinear constitutive relation (DNCR) method as a promising approach for expeditious physical modeling of non-equilibrium rarefied flows. To enhance the generalization capabilities of the DNCR method, this study proposes a deep convolutional neural network model (DNCR-CNN) based on data-driven nonlinear constitutive relations, integrated with free-form deformation (FFD). Employing FFD technology, a series of hypersonic geometric shapes are generated for model training, and a multi-task learning-based deep convolutional neural network model is subsequently trained. The prediction of the hypersonic geometric shapes test set is carried out, and the results of the model prediction are substituted in the conservation equation for the iterative solution, thereby enhancing the DNCR method's generalization performance for varying geometric shapes. Upon conducting a comparative analysis of the outcomes obtained from DNCR, Navier–Stokes (NS), and unified gas kinetic scheme (UGKS), it is revealed that the DNCR method can maintain computational resource levels equivalent to those of the NS equation while achieving a level of accuracy comparable to UGKS under diverse geometric shapes and grid resolutions. The enhancements in usability render the DNCR method a potent tool for addressing the challenges posed by rarefied gas, thereby expanding its applicability within the field.

Adaptive Neural Network Global Fractional Order Fast Terminal Sliding Mode Model-Free Intelligent PID Control for Hypersonic Vehicle’s Ground Thermal Environment

In this paper, an adaptive neural network global fractional order fast terminal sliding mode model-free intelligent PID control strategy (termed as TDE-ANNGFOFTSMC-MFIPIDC) is proposed for the hypersonic vehicle ground thermal environment simulation test device (GTESTD). Firstly, the mathematical model of the GTESTD is transformed into an ultra-local model to ensure that the control strategy design process does not rely on the potentially inaccurate dynamic GTESTD model. Meanwhile, time delay estimation (TDE) is employed to estimate the unknown terms of the ultra-local model. Next, a global fractional-order fast terminal sliding mode surface (GFOFTSMS) is introduced to effectively reduce the estimation error generated by TDE. It also eliminates arrival time, accelerates the convergence speed of the sliding phase, guarantees finite time arrival, avoids the singularity phenomenon, and bolsters robustness. Then, as the upper bound of the disturbance error is unknown, an adaptive neural network (ANN) control is designed to approximate the upper bound of the estimation error closely and mitigate the chattering phenomenon. Furthermore, the stability of the control system and the convergence time are proven by the Lyapunov stability theorem and are calculated, respectively. Finally, simulation results are conducted to validate the efficacy of the proposed control strategy.

Adaptive Model Reduction of High-Order Solutions of Compressible Flows via Optimal Transport

The solution of conservation laws with parametrised shock waves presents challenges for both high-order numerical methods and model reduction techniques. We introduce an r-adaptivity scheme based on optimal transport and apply it to develop reduced order models for compressible flows. The optimal transport theory allows us to compute high-order r-adaptive meshes from a starting reference mesh by solving the Monge–Ampère equation. A high-order discretization of the conservation laws enables high-order solutions to be computed on the resulting r-adaptive meshes. Furthermore, the Monge–Ampère solutions contain mappings that are used to reduce the spatial locality of the resulting solutions and make them more amenable to model reduction. We use a non-intrusive model reduction method to construct reduced order models of both the mesh and the solution. The procedure is demonstrated on three supersonic and hypersonic test cases, with the hybridisable discontinuous Galerkin method being used as the full order model.

Method of applying temperature-sensitive paint in hypersonic test with strong combustion radiation

Method of applying temperature-sensitive paint in hypersonic test with strong combustion radiation

A new aerodynamic identification technology for short-time hypersonic wind tunnels while considering inertial force interference

When aerodynamic force tests are performed in hypersonic wind tunnels, the test model in the force measurement system (FMS) will be subjected to transient impact and induced transient vibration due to the high-speed airflow generated during start-up. The output signal of the FMS will contain the real aerodynamic signal and the inertial force signal caused by model vibration. The inertial force signal completely overcomes the aerodynamic signal, strongly influencing the test accuracy of the aerodynamic force and even making the aerodynamic force signal completely unusable. Therefore, this study develops a novel intelligent aerodynamic identification method based on deep residual shrinkage network (DRSN) deep learning technology and applies this method to aerodynamic force tests in hypersonic wind tunnels. This method can achieve intelligent identification and filtering of inertial force interference signals and instrument noise signals, and improve the test accuracy of aerodynamic forces. Also, this study performed a numerical simulation experiment and a test-bed experiment; verified the results through hypersonic wind tunnel test signals; and compared the identification results of the DRSN network method with the identification results of the traditional CNN network method, the mean method and the FFT method. From comparison results, the aerodynamic force identification accuracy of the DRSN network method proposed in this paper is shown to be nearly ideal and always remains above 95%. Thus, the proposed method effectively reduces the interference of instrument noise and inertial force, markedly improving the accuracy and reliability of the hypersonic wind tunnel aerodynamic force test.

High-speed penetration of ogive-nose projectiles into thick concrete targets: Tests and a projectile nose evolution model

The majority of the projectiles used in the hypersonic penetration study are solid flat-nosed cylindrical projectiles with a diameter of less than 20 mm. This study aims to fill the gap in the experimental and analytical study of the evolution of the nose shape of larger hollow projectiles under hypersonic penetration. In the hypersonic penetration test, eight ogive-nose AerMet100 steel projectiles with a diameter of 40 mm were launched to hit concrete targets with impact velocities that ranged from 1351 to 1877 m/s. Severe erosion of the projectiles was observed during high-speed penetration of heterogeneous targets, and apparent localized mushrooming occurred in the front nose of recovered projectiles. By examining the damage to projectiles, a linear relationship was found between the relative length reduction rate and the initial kinetic energy of projectiles in different penetration tests. Furthermore, microscopic analysis revealed the forming mechanism of the localized mushrooming phenomenon for eroding penetration, i.e., material spall erosion abrasion mechanism, material flow and redistribution abrasion mechanism and localized radial upsetting deformation mechanism. Finally, a model of high-speed penetration that included erosion was established on the basis of a model of the evolution of the projectile nose that considers radial upsetting; the model was validated by test data from the literature and the present study. Depending upon the impact velocity, v0, the projectile nose may behave as undistorted, radially distorted or hemispherical. Due to the effects of abrasion of the projectile and enhancement of radial upsetting on the duration and amplitude of the secondary rising segment in the pulse shape of projectile deceleration, the predicted DOP had an upper limit.

Installation error calibration for MEMS angular measurement system in hypersonic wind tunnel

The accuracy of model attitude measurement has an important impact on wind tunnel test results. Microelectromechanical System Inertial Measurement Unit (MEMS IMU) provides a feasible way to measure model attitudes with high accuracy. However, the installation error between MEMS IMU coordinate system and the body coordinate system of test models can make the accuracy of the model attitude measurement decrease. In wind tunnel tests, the installation error depends on the relationship between the IMU and the model mechanism before tests. Therefore, in-field calibration in wind tunnel tests is necessary to reduce installation errors. To improve attitude measurement accuracy, the least squares quaternion calibration method based on MEMS IMU and six-position calibration procedure are proposed. High-precision three-axis turntable tests are performed. The pitch accuracy after calibration is higher than that before calibration in the angle of attack sweeping tests. The Root-Mean-Square Errors (RMSE) in the roll and yaw are within 0.01°, which are smaller than those before calibration. In the roll sweeping tests, RMSE of three attitude angles decrease significantly. In hypersonic wind tunnel tests, the pitch errors before and after calibration are within 0.05° and 0.02° in the angle of attack sweeping tests without wind. In five angle of attack sweeping tests with wind, the deviation between the mean of the pitch and the pitch after the elastic angle correction is within 0.03° and the standard deviation of five tests is within 0.01°. The proposed method is confirmed to enhance the accuracy of attitude measurement effectively, which is convenient for engineering applications.

Stagnation heat flux estimation in spherically blunt axisymmetric hypersonic models

Hypersonic flows have high heat transfer rates, and their management is essential to avoid detrimental effects. Since accurate prediction and measurement of heat flux in hypersonic test facilities are complicated, heat flux at the stagnation point is mostly estimated using Fay and Riddell formulation with Newtonian tangential velocity gradient approximation. Although it is relatively accurate and reliable, some errors creep in due to incompetent modelling of the tangential velocity gradient. This article studies the applicability of Olivier's tangential velocity gradient formulation for a sphere in the estimation of stagnation heat flux for spherically blunt axisymmetric hypersonic models. Oliver’s estimation accurately models the tangential velocity gradient of spherically blunt axisymmetric hypersonic models as the heat flux estimates deviated only by approx. 2%–4% from the measured heat flux. A simplified model for tangential velocity gradient using Shock Standoff Distance and density ratio is also derived and tested for accuracy.

Numerical tripping of high-speed turbulent boundary layers

The influence of turbulence inflow generation on direct numerical simulations (DNS) of high-speed turbulent boundary layers at Mach numbers of 2 and 5.84 is investigated. Two main classes of inflow conditions are considered, based on the recycling/rescaling (RR) and the digital filtering (DF) approach, along with suitably modified versions. A series of DNS using very long streamwise domains is first carried out to provide reliable data for the subsequent investigation. A set of diagnostic parameters is then selected to verify achievement of an equilibrium state, and correlation laws for those quantities are obtained based on benchmark cases. Simulations using shorter domains, with extent comparable with that used in the current literature, are then carried out and compared with the benchmark data. Significant deviations from equilibrium conditions are found, to a different extent for the various flow properties, and depending on the inflow turbulence seeding. We find that the RR method yields superior performance in the evaluation of the inner-scaled wall pressure fluctuations and the turbulent shear stress. DF methods instead yield quicker adjustment and better accuracy in the prediction of wall friction and of the streamwise Reynolds stress in supersonic cases. Unrealistically high values of the wall pressure variance are obtained by the baseline DF method, while the proposed DF alternatives recover a closer agreement with respect to the benchmark. The hypersonic test case highlights that similar distribution of wall friction and heat transfer are obtained by both RR and DF baseline methods.Graphical abstract

The Texas A&M University Hypervelocity Impact Laboratory: A modern aeroballistic range facility.

Novel engineering materials and structures are increasingly designed for use in severe environments involving extreme transient variations in temperature and loading rates, chemically reactive flows, and other conditions. The Texas A&M University Hypervelocity Impact Laboratory (HVIL) enables unique ultrahigh-rate materials characterization, testing, and modeling capabilities by tightly integrating expertise in high-rate materials behavior, computational and polymer chemistry, and multi-physics multiscale numerical algorithm development, validation, and implementation. The HVIL provides a high-throughput test bed for development and tailoring of novel materials and structures to mitigate hypervelocity impacts (HVIs). A conventional, 12.7mm, smooth bore, two-stage light gas gun (2SLGG) is being used as the aeroballistic range launcher to accelerate single and simultaneously launched projectiles to velocities in the range 1.5-7.0 km/s. The aeroballistic range is combined with conventional and innovative experimental, diagnostic, and modeling capabilities to create a unique HVI and hypersonic test bed. Ultrahigh-speed imaging (10M fps), ultrahigh-speed schlieren imaging, multi-angle imaging, digital particle tracking, flash x-ray radiography, nondestructive/destructive inspection, optical and scanning electron microscopy, and other techniques are being used to characterize HVIs and study interactions between hypersonic projectiles and suspended aerosolized particles. Additionally, an overview of 65 2SLGG facilities operational worldwide since 1990 is provided, which is the most comprehensive survey published to date. The HVIL aims to (i) couple recent theoretical developments in shock physics with advances in numerical methods to perform HVI risk assessments of materials and structures, (ii) characterize environmental effects (water, ice, dust, etc.) on hypersonic vehicles, and (iii) address key high-rate materials and hypersonics research problems.

Numerical Assessment of Camphor Ablation Flight Relevance in Hypersonic Wind-Tunnel Testing

Designing thermal protection systems (TPS) for ballistic reentry capsules requires both reliable numerical tools and experimental facilities to evaluate surface ablation and its shape change. Low-temperature ablators, such as camphor, can be employed in low-enthalpy wind tunnels to overcome limitations of standard hypersonic experimental facilities, which cannot induce the actual TPS ablation due to limits in inflow total enthalpy, velocity, or test time duration. However, the flight relevance of such an experiment needs to be accurately assessed, as both the capsule dimension and the ablative material during the test differ from the flying ones. The main goal of this work is to perform numerical simulations in order to assess the flight relevance of the ground tests. Computational fluid dynamics (CFD) simulations have been performed for both low-enthalpy wind-tunnel conditions for a camphor scaled-down capsule and actual reentry flight conditions, assuming a carbon-based heat shield. Moreover, numerical analyses have been improved by coupling the CFD solver with a material response code named PATO, which models the heating transient process. The obtained results show how the CFD/material response coupling has a more important effect for the on-ground testing conditions considered, resulting in a more favorable agreement with the available experimental data. Finally, the shape change relevance of the analyzed low-temperature ablator has been verified by providing a similitude criterion between the on-ground and in-flight capsule recession. In particular, a few seconds of exposure are sufficient to generate the same shape change that is expected in flight.

Joule heated hot wall models in hypersonic wind tunnel facilities for energy deposition studies

Joule heating is a useful experimental technique for ground based hypersonic testing as it allows replication of the governing wall-to-freestream temperature ratio and thus better simulation of the aerothermodynamic environment of high-speed flight vehicles. This paper presents design considerations for using this hot-wall method for hypersonic wind tunnel models. A numerical analysis of material choice and thickness for creating rapid uniform high-temperature model conditions is presented. The influence of the method and location of energisation and structural requirements for the integration of joule heating in a hypersonic wind tunnel model is also examined numerically. Experimental results of a bench-top study incorporating the design elements for joule heating for studies analogues to thermal energy deposition are also presented. To replicate thermal energy deposition, graphite plates with a 1.0 or 2.5mm thickness were heated via joule heating to a base level of either 30W or 100W. After 120s of heating the plates were then exposed to a thermal pulse of either 550, 1200 or 1300W. The thermal-structural response of the plate was experimentally recorded with a measurement system incorporating infrared thermal cameras, single and two-colour pyrometers, a load cell and a 520 nm light Digital Image Correlation system suitable for use on hot wall models. Results from this bench-top study demonstrate a proof-of-concept for using joule heating for analogue thermal deposition studies. Temperature rises of ΔT on the order of 500 - 1000K were achieved on models with base temperatures of 450 - 600K achieved with this technique. Plate deflections as high as 3.2 - 4.1mm were recorded for 1.0mm plates with both stress and strain of the heated models resolved experimentally for base heating and pulsed energy conditions.

Computation of hypersonic viscous flows with the thermally perfect gas model using a discontinuous Galerkin method

AbstractThis article is concerned with the numerical simulation of high‐speed thermally perfect flows using high‐order discontinuous Galerkin (DG) methods. In these methods, the polynomial approximation of the solution is susceptible to spurious oscillations, especially in high‐gradient regions. The introduction of variable thermodynamic properties, important for hypersonic flows, can exacerbate this issue. Given the general lack of satisfactory high‐order DG simulations of these types of flows, our primary objective is to demonstrate the success of the formulation proposed here in computing flows in this regime. Secondary goals are to (a) highlight that DG schemes are significantly less sensitive than finite‐volume methods to mesh‐shock misalignment in the given setting and (b) evaluate the sensitivity of surface heating predictions to various types of no‐slip isothermal wall boundary conditions. To handle strong shocks prevalent in this flow regime, intra‐element solution variations are used to detect discontinuities and smooth artificial viscosity is employed for stabilization. We discuss the algorithmic developments for extending a standard DG discretization of the compressible Navier–Stokes equations with the calorically perfect gas model to the thermally perfect gas model. We apply the resulting DG formulation to several hypersonic test cases, including inviscid flow over a cylinder and viscous flow over a double cone. Where appropriate, our results are benchmarked against those obtained with state‐of‐the‐art finite‐volume hypersonic flow solvers. The final case entails high‐speed particle‐laden flow over a blunt body. The results of this test illuminate how the consideration of variable thermodynamic properties influences the prediction of the particle phase.