Hypersonic Wind Tunnel Test Research Articles

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

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.

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.

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.

A systematic approach for horizontal and vertical scale up of sintered Ultra-High Temperature Ceramic Matrix Composites for aerospace – Advances and perspectives

Ultra-High Temperature Ceramic Matrix Composites (UHTCMCs) are the next generation composites developed for application in the harsh environment of aerospace. Qualification of these materials requires the manufacturing of demonstrators for testing in relevant environments, which in turn, requires to scale them up. This paper illustrates the systematic approach adopted for the scale-up of UHTCMCs based on carbon fibre reinforced zirconium diboride plus silicon carbide from laboratory to industrial scale dimensions. The scale-up process covered a period of three years and concerned an increase in diameter of ∼10 times (from 40 to 400 mm), and in thickness of ∼30 times (from 5 to 160 mm). Small-scale products were consolidated by hot pressing at ISTEC whilst larger samples were consolidated by an industrial spark plasma sintering facility (NanokerResearch, Spain). After each scale-up step, reproducibility of composition, microstructure and properties was carefully checked. Thanks to the homogenous fibre distribution and sapient dosing of secondary phases, composites with different diameters and thickness were successfully consolidated by both techniques with little adjustment of sintering parameters up to the maximum size of available industrial furnaces . The large discs produced allowed production of 170 mm long bars for tensile testing and large tiles for hypersonic wind tunnel tests . Thick samples were useful to machine complex shapes such as screws and nuts, vertical bars for characterization of properties along the composite thickness and nozzle demonstrators. • ZrB 2 -Cf composites were scaled-up 10 times in diameter, 30 times in thickness. • A manufact of UHTCMC material weighting 11 kg was produced for the first time. • The same mechanisms were efficient for scale-up in diameter and thickness. • Reproducible microstructure-properties were assessed from lab to industrial scale. • Large components such as plates, nozzle inserts, screw and nuts were fabricated.

A New Hypersonic Wind Tunnel Force Measurement System to Reduce Additional Bending Moment and Avoid Time-Varying Stiffness.

In order to improve traditional hypersonic wind tunnel airframe/propulsion integrated aerodynamic testing technology for hypersonic vehicles, a new force measurement system called the aerodynamic force measuring support (AFMS) was designed. The AFMS effectively overcomes the defect that the traditional internal box-balance occupies a large amount of internal space in the aircraft test model, which makes the airframe/propulsion integrated aerodynamic test more difficult. The AFMS also alleviates the interference of the additional bending moment caused by the non-coincidence between the calibration center of traditional external box-balance and the gravity center of the aircraft test model, innovatively designing a convex structure in the joint part of the force measuring system. Furthermore, the AFMS effectively overcomes the time-varying stiffness of joints caused by test model vibration in hypersonic wind tunnel testing, which eventually leads to test errors. Compared with the traditional box-balance, the AFMS proposed in this study has sufficient design space. This ensures more thorough aerodynamic decomposition of the AFMS and less interference between channels, whilst also having the advantages of the large support stiffness of traditional box-balance. Thus, the AFMS provides a new technical path for airframe/propulsion integrated aerodynamic testing of air-breathing hypersonic vehicles in a hypersonic wind tunnel.

A New Airframe/Propulsion-Integrated Aerodynamic Testing Technology in Hypersonic Wind Tunnel

This article presents a new type of multicomponent force measurement system to measure lift force, drag force, and pitching moment in hypersonic wind tunnel, which called the suspension force-measuring system (SFMS). The SFMS does not occupy the installation space of the scramjet engine and other key functional components in the air-breathing hypersonic vehicle model. It solves the problem that the conventional aerodynamic force measurement technology occupies too much internal space of the hypersonic vehicle test model and dramatically reduces the design difficulty of scramjet propulsion system in the test model. The SFMS adopts the spatial multi-point distribution support mode for the test model. Therefore, it has an excellent stiffness characteristic, meeting the requirement of support rigidity of wind tunnel force measurements test for large-scale and heavy load test model. Then the designed SFMS is constructed and calibrated. The static calibration results show that each test channel of the SFMS has a good sensitivity of 0.2502 mV/N for drag force test channel, 0.1557 mV/N for lift force test channel, 0.2484 mV/Nm for pitching moment test channel, and the max interference error among the SFMS test channels is less than 2.3&#x0025;. A series of aerodynamic force experiments have been carried out in the <inline-formula> <tex-math notation="LaTeX">$\emptyset 600~mm$ </tex-math></inline-formula> impulse combustion wind tunnel, and the output data of each test channel well reflected the change process of the total pressure in the test section of wind tunnel. Hence, the proposed device can meet the force test requirements of the hypersonic wind tunnel test.

Heat Transfer and Recovery Factor of Aerodynamic Heating on a Flared Cone

This paper investigates aerodynamic heating on a flared cone in a Mach 6 wind tunnel. Two data processing methods to calculate the Stanton number and recovery factor simultaneously are presented, using the experimentally measured wall temperature in a hypersonic wind-tunnel test. The first method is the linear fitting of the measured temperature and calculated heat flux. The second method is to solve an inverse problem of a one-dimensional transient heat conduction equation with the convection boundary condition. Distributions of heat flux, Stanton number, and recovery factor for the flared cone with a smooth wall or wavy wall are obtained. Compared with the smooth wall, the wavy wall can remove the local high Stanton number region before the boundary-layer transition. In addition, the wavy wall can reduce the growth rate of Stanton number in the boundary-layer transition and delay the transition process.

A stereo photogrammetry system for model position and attitude measurement in hypersonic wind tunnel testing.

In this paper, a measurement system based on stereo photogrammetry is proposed for position and attitude measurements of test models in hypersonic wind tunnel testing. Several adaptations were made to an existing stereo photogrammetry system to promote its routine use as a primary wind tunnel measurement system. First, cameras were installed in the wind tunnel test section to decrease the measurement distance and increase the base length, which helped improve the measurement accuracy. To deal with camera vibration caused by the freestream, a camera orientation method is proposed to estimate the exterior parameters of each frame relative to the tunnel coordinate system. Second, a scheme to determine the model coordinate system is proposed, which does not require placement of the targets at known locations of the model. Third, to solve the recognition problem of motion-blurred target images, a network was adopted for stereo deblurring and a stereo blur dataset simulating the wind tunnel test environment was generated for network training. To develop the novel stereo photogrammetry system, experiments were conducted in both the laboratory and a hypersonic wind tunnel at the China Aerodynamics Research and Development Center. The experimental results suggest that the proposed measurement system is practicable for position and attitude measurements of test models in hypersonic wind tunnel testing.

A review of theories and methods for hypersonic nozzle design

In the field of hypersonic flight technology, especially the study of high-enthalpy flows, the hypersonic wind tunnel test, at present, is still the most reliable and convenient research method. For the development of hypersonic wind tunnels, one of the most important performance indicators is the uniform flow quality, which depends on the theories and methods of hypersonic nozzle design, and is the core technique of wind tunnel design. For the design of a two-dimensional axisymmetric contour nozzle, this paper first reviews the main theories and traditional design methods that have played an important role in the design of hypersonic nozzles, including the theoretical methods, the approximate methods, and the correction methods for both. Then, due to the high-temperature gas effects, the difficulties and problems faced in the design of high-enthalpy nozzles are analyzed in detail. In terms of the changes in the physical properties of the test gases, the development of high-temperature boundary-layer, and the effects of non-equilibrium processes, this paper also reviews the domestic and foreign research progresses in the design of hypersonic and high-enthalpy nozzles. Finally, the development of theory and method for designing the high-enthalpy nozzle is prospected, and we expect it to promote the development of hypersonic nozzle design technology in China.

Visual three-dimensional measurement of aircraft flutter in hypersonic wind tunnel tests

Visual three-dimensional measurement of aircraft flutter in hypersonic wind tunnel tests

Effect of Insulations and Coatings on Hypersonic Heat Flux Reconstruction

The present study investigates the effects of material discontinuities and thermal resistances on the temperature reconstruction and heat flux evaluation via an inverse heat transfer method in hypersonic wind-tunnel tests. Thin coating layers and adhesive interfaces support a localized multidimensional temperature perturbation that alters the near-surface field. A technique is developed for the real-time evaluation of the heat transfer based on approximating the Green’s function of the heat conduction equation in bounded complex domains and resolving the heat flux on a B-spline basis, leading to an efficient and robust algorithm. Evaluating the Green’s function for three-dimensional problems with material discontinuities is generally difficult. A new approach is developed in which the Green’s function is evaluated independently for the regular partitions of the domain, and appropriate connection conditions are derived to determine the full “system” Green’s function. This paper describes the verification, validation, and application of the approach to the heat flux reconstruction for a nonaxisymmetric flat-faced article with coating and potting interfaces. The article was tested at Mach 6 in the Virginia Tech Hypersonic wind tunnel using four temperature sensors arranged in a cross pattern on the flat face. Good agreement was obtained between computational fluid dynamics predictions and the heat flux deduced from the experiments using the new data analysis method.

The effect of concave roughness on the hypersonic flat-plate boundary-layer transition

In order to reveal the effect of aircraft surface concave roughness on the transition, several hypersonic quiet wind tunnel tests were carried out on the flat-plate model with independent concave roughness on the surface. The experimental Mach number was 6.5 and the total temperature was 430 K. Using infrared thermal imaging technology, a general rule of the effect of concave roughness on the boundary layer transition was obtained. The flow visualization results and the PCB pressure measurement results showed that the concave roughness promotes the transition by increasing the growth of the first mode wave rather than the growth of the second mode wave.

Gamma and infrared novel methodologies in Aerospace re-entry: γ-rays crystal efficiency by GEANT4 for TPS material recession assessment and simultaneous dual color infrared temperature determination

The re-entry of a vehicle into Earth's atmosphere either from near-Earth orbit or from interplanetary trajectories is the most critical phases of a space mission for materials, since space capsules and spacecraft surfaces are brought to relatively high temperatures (>2000 °C) and are subject to the high thermal fluxes. The most complex requirements of the on-ground CIRA hypersonic Plasma Wind Tunnel (PWT) test campaigns are to measure and monitor the high heat fluxes, the correlated surface Recession Rate and Surface Temperatures affecting the behavior of Thermal Protection System (TPS) materials which makes up the surface and sub-surface components of the spacecraft's outer layer.For these tests, the recession rate will be obtained by Surface Layer Implantation (SLI), monitored by the gamma ray emission from 7Be implanted into the test material, and the surface temperature by Dual Color free emissivity Infrared Thermography (DCIT). So the first purpose of the paper, in the recession rate determination, is to present the γ-ray detection efficiency of a LaBr3(Ce) crystal by GEANT4 simulation using the radioactive ions, to be implanted by means of the CIRCE accelerator, and the use of γ-rays to monitor their removal during the ablation process. At the same time, for temperature determination, is to evaluate the possibility of carrying out free emissivity temperature measurements in unsteady conditions.On the basis of the results, possible validation tests are outlined, through the use of the intense radioactive 7Be beam available at the 3MV Pelletron tandem accelerator of the CIRCE Laboratory and the GHIBLI-PWT facility at CIRA, to present the combined novel, contactless and non-intrusive diagnostics for the simultaneous recession rate and temperature determinations.

Pyrotechnic approach to space debris destruction: From thermal modeling to hypersonic wind tunnel tests

The problem of the space debris is an important aspect of the on-ground casualty risk reduction. Some articles, like fuel tanks made of titanium alloy, survive during the atmospheric post-mission re-entry and impact the Earth's surface, thus posing a serious hazard. A pyrotechnic device is proposed to bring an additional heating of space debris at a certain altitude promoting its destruction and fragmentation into smaller parts during re-entry. 3D thermal modeling of the non-stationary thermal wave propagation through the system ‘pyrotechnic device - fuel tank wall’ has been performed alongside with extensive experimental testing. The pyrotechnic composition (namely, thermite), the design of the pyrotechnic device, the conjunction between the device, and the target plate surface were optimized during modeling. The first level of experimental verification included lab-tests with laser heating. At the second level, large-scale tests were conducted within the hypersonic wind tunnel to achieve the conditions closest to the atmospheric re-entry. The presented results reveal the usefulness of pyrotechnics for space applications; and the performed thermal engineering approach can be extended to materials other than titanium to decrease the on-ground casualty risk of space debris.

Numerical study of total temperature effect on hypersonic boundary layer transition

The total temperature in the hypersonic wind tunnel test can be significantly different from that in real flight conditions, and this leads to a large discrepancy in the measurement of the boundary layer transition between ground experiment measurements and flight tests. Even at the same Mach number and Reynolds number, different wind tunnels may yield different transition data for the same model due to the total temperature effect. In this paper, the boundary layer transition on a 7° half-angle sharp cone (at a 0° angle of attack) with four freestream total temperatures is investigated using both the simulations of a local correlation-based transition model and linear stability analysis. The results show that as the freestream total temperature increases, the starting point of the transition on the sharp cone gradually moves backward and the length of the transition region decreases. The N factor of the unstable wave gradually decreases with increasing freestream total temperature, causing the transition onset to move backward. The total temperature effects on boundary transition as determined by both methods of analysis were in good agreement.

Hypersonic Heat Flux Reconstruction with Distributed Temperature Sensors

This paper presents a new method for the multidimensional reconstruction of the heat flux via distributed surface temperature measurements and its validation using hypersonic wind tunnel tests. The first innovation is the formulation of the reconstruction in terms of the Green’s function, approximated using a weak Galerkin technique together with the method of variation of parameters. When compared with previous research, the method does not require nullification of the first derivative of the polynomial basis at boundaries of the second kind, and it applies seamlessly to complex geometries. The second innovation is the use of homotopy to reduce the nonlinear conduction at high temperature to a series of linear subproblems. Our main contribution is the determination of the conditions for the quadrature formulas to yield a convergent homotopy sequence. The third innovation is the use of a multidimensional B-spline basis to reconstruct the heat flux on curved surfaces and in time. The inverse problem projected in the B-spline basis is well-posed and does not require regularization parameters. The fourth innovation is the introduction of a piecewise Green’s function on partitioned multimaterial domains by imposing continuity of heat flux and temperature. The partitioned approach improves the accuracy of the reconstruction in the linear and nonlinear regimes.

Accurate attitude estimation of HB2 standard model based on QNCF in hypersonic wind tunnel test

The accuracy of the HB2 standard model attitude measurement has an important impact on the hypersonic wind tunnel data assessment. The limited size of the model and the existence of external vibrations make it challenging to obtain real-time reliable attitude measurement. To reduce the influence of attitude errors on test results, this paper proposes a Quaternion Nonlinear Complementary Filter (QNCF) attitude determination algorithm based on Microelectromechanical Inertial Measurement Unit (MEMS-IMU). Firstly, the threshold-based PI control strategy is adopted to eliminate noise effect according to the Acceleration Magnitude Detector (AMD). Then, the flexible quaternion method is updated to carry out attitude estimation which is operational and easy to be embedded in the Field Programmable Gate Array (FPGA). Finally, a high-precision three-axis turntable test and a hypersonic wind tunnel test are performed. The results show that the pitch-roll attitude errors are within 0.05° and 0.08° in the high-precision three-axis turntable test in a calculation time of 100 s respectively, and the attitude error is within 0.3° after the elastic angle correction in the hypersonic wind tunnel test. The proposed method can provide accurate real-time attitude reference for the analysis of the actual movement of the model, exhibiting certain engineering application value with robustness and simplicity.

Suspension force measuring system for hypersonic wind tunnel test: Design and tests

Suspension force measuring system is a new type of measuring system for dynamometric test of hypersonic aircraft wind tunnel. The system alters the measuring solution of internal balance placed in the model chamber for wind tunnel test in the past to the new design without internal balance, which dramatically reduces design difficulty of aircraft model with engine. The processing, installation and calibration test of this system were carried out, it turns out that the system performs well in repeatability and linearity in various measure components, line repeatability of calibration data point by least squares fit is good, and maximum interference between components is less than 2.3%. Dynamic test was also carried out, frequency spectrum shows the first-order frequency with the greatest contribution to the vibration of the system reach to 137 Hz, dynamic measurement capability performs well. Finally, the system is used in hypersonic wind tunnel test and the stress state of the model under the action of high-speed airflow was correctly reflected.

Development of high temperature material for cellular ceramic regenerative storage heater

Ceramic Regenerative Storage Heater (RSH) is the preferable solution to provide clean hot air in Hypersonic Wind Tunnel (HWT) and Scramjet Test facilities to prevent the liquefaction of air during expansion in the high Mach number nozzles. Selection of the heater storage material for such operational requirements is a challenging task. As per the theoretical background; alumina, zirconia or mullite compositions have the capability to withstand at such high operating conditions. In the current study, a series of experiments were carried out to estimate the thermal characteristics of the material. This paper highlights the simplified approach by combining theoretical and experimental methods for economic (reduced time and cost) selection process of high temperature material. The desired material must possess good thermal shock resistance and low dust generation characteristics. Al2O3 and SiO2 based matrix composition fused in a transcrystalline phase exhibits significant improvement on the overall material integrity. Several Al2O3 and SiO2 matrix compositions were examined and the final matrix composite ratio of 90% of Al2O3 and 10% of SiO2 is considered as the desired candidate material. The ceramic matrix was sintered at a temperature of 2123 K, under the specific pressure of 1400 kg/cm2 for a period of 2 h, at a heating rate of 3.5 °C/min. The results of the thermal shock tests on various samples are presented. To improve the anti-dust characteristics, the final samples were coated with fine particles (5 µm) of Al2O3 and SiO2 matrix. Based on the experimental evaluation, Al2O3 and SiO2 fused matrix composition found to be meeting all the requirements of high temperature storage material.