Shock Tunnel Research Articles

Converging near-elliptic shock waves

Abstract

Experimental study of boundary-layer transition on a 7° sharp cone in shock tunnel

Experimental study of boundary-layer transition on a 7° sharp cone in shock tunnel

Criteria for hypersonic airbreathing propulsion and its experimental verification

Hypersonic airbreathing propulsion is one of the top techniques for future aerospace flight, but there are still no practical engines after seventy years' development. Two critical issues are identified to be the barriers for the ramjet-based engine that has been taken as the most potential concept of the hypersonic propulsion for decades. One issue is the upstream-traveling shock wave that develops from spontaneous waves resulting from continuous heat releases in combustors and can induce unsteady combustion that may lead to engine surging during scramjet engine operation. The other is the scramjet combustion mode that cannot satisfy thrust needs of hypersonic vehicles since its thermos-efficiency decreases as the flight Mach number increases. The two criteria are proposed for the ramjet-based hypersonic propulsion to identify combustion modes and avoid thermal choking. A standing oblique detonation ramjet (Sodramjet) engine concept is proposed based on the criteria by replacing diffusive combustion with an oblique detonation that is a unique pressure-gain phenomenon in nature. The Sodramjet engine model is developed with several flow control techniques, and tested successfully with the hypersonic flight-duplicated shock tunnel. The experimental data show that the Sodramjet engine model works steadily, and an oblique detonation can be made stationary in the engine combustor and is controllable. This research demonstrates the Sodramjet engine is a promising concept and can be operated stably with high thermal efficiency at hypersonic flow conditions.

Single image molecular tagging velocimetry

A single-image nitric-oxide molecular tagging velocimetry (NO-MTV) is reported and employed in a real air driven hypersonic shock tunnel and provided velocities of 3240 ± 170(5.2%) and 3030 ± 160(5.3%) m , insignificantly different from previous results of 3037 ± 98(3.2%) obtained by an ordinary multi-image MTV technique. The proposed methodology relies on an one-dimensional analytical description of the spatial intensity profile registered by a single MTV image.

P.261 Air pollution and bipolar disorder: is there an association between manic episodes and PM10 exposure?

A blunt bi-conic aluminum model, integrated with a three component accelerometer force balance system, is tested in the IITB-Shock Tunnel at various angle of inclinations with an intention to measure aerodynamic coefficients. Initially, Genetic Algorithm (GA) is employed to deduce the orthogonal inputs and their responses from distributed point loads applied on the test model during calibration experiments. Time histories of these loads and their responses are then used to train the architecture of Adaptive Neuro Fuzzy Inference System, ANFIS. This ANFIS architecture is further employed for force prediction from the acquired acceleration responses during shock tunnel experiments. Testing of the experimental acceleration responses of 0° and 10° angle of attack experiments showed encouraging agreement for recovered aerodynamic coefficients with the accelerometer balance theory based predictions. These results clearly showed the necessity to consider multiple point loads and their acceleration responses for training the soft computing algorithm or calibration of the force balance. Further, use of only one point force and its responses, for training ANFIS, is not only seen to have discrepancy in prediction but also is noted to be non-unique due to choice of the loading point. Moreover, it is recommended to choose the loading point near the center of pressure for the tested experimental conditions to incur lower discrepancy in prediction using single point loading data for ANFIS training.

Attenuation of pulsation and oscillation using a disk at mid-section of spiked blunt body

The two most important modes of flow instability are observed on a flat-faced cylinder fitted with a spike, namely, pulsation and oscillation. A new technique is proposed in this article to attenuate both pressure and aerothermal fluctuations over the cylinder by using a circular disk at the mid-length of the sharp-tipped spike, which splits the separation region in front of the cylinder. Shock tunnel experiments are carried out at a free stream Mach number of M∞ = 5.7 and a Reynolds number of RD = 0.5 × 106 (based on the cylinder diameter, D). Aeroheating is measured using platinum thin film sensors painted over the front face of the cylinder. It is shown that the flow pulsation can be totally mitigated for the ratio of spike length to diameter (L/D = 1.0) by using a disk of diameter d/D = 28% at the mid-section of the spike, and the oscillation flow is suppressed using the same disk diameter with L/D = 1.5. This phenomenon is well evident from the measured heat transfer and schlieren data. Besides, a similar analysis showed that the flow fluctuation over an aerospike mounted on a hemispherical blunt body can be diminished with the use of double-aerospike, which, in turn, decreases the heat flux peak in the proximity of the reattachment. This is an effective means of flow fluctuation mitigation and presents to be a good thermal protection for the fore body.

Intelligent Force-Measurement System Use in Shock Tunnel.

The inertial vibration of the force measurement system (FMS) has a large influence on the force measuring result of aircraft, especially on some tests carried out in high-enthalpy impulse facilities, such as in a shock tunnel. When force tests are conducted in a shock tunnel, the low-frequency vibrations of the FMS and its motion cannot be addressed through digital filtering because of the inertial forces, which are caused by the impact flow during the starting process of the shock tunnel. Therefore, this paper focuses on the dynamic characteristics of the performance of the FMS. A new method—i.e., deep-learning-based single-vector dynamic self-calibration (DL-based SV-DSC) of an impulse FMS, is proposed to increase the accuracy of aerodynamic force measurements in a shock tunnel. A deep-learning technique is used to train the dynamic model of the FMS in this study. Convolutional neural networks with a simple structure are applied to describe the dynamic modeling so that the low-frequency vibration signals are eliminated from the test results of the shock tunnel. By validation of the force test results measured in a shock tunnel, the current trained model can realize intelligent processing of the balance signals of the FMS. Based on this new method of dynamic calibration, the reliability and accuracy of force data processing are well verified.

Investigation of the Flow In Cold Condition at the Exit of a Supersonic Combustor Test Bench

For studies of hypersonic flows and supersonic combustion in ground test facilities, three devices can be used as ram accelerators, shock tunnels and supersonic combustor test benches. These devices can reproduce, on the ground, similar conditions to those in real flight at a certain altitude and speed. In the case of the supersonic combustor test bench (SCTB), it simulates the same flow conditions inside the combustor of a scramjet. The SCTB consists basically of a combustion chamber or vitiated air generator unit, where the air is heated, and a nozzle, where the air is accelerated to the desired test speed. The supersonic combustor to be tested is directly coupled to the nozzle exit of the SCTB. Ultimately, it was necessary to use a transition piece to connect the nozzle to the combustor to be tested, because the nozzle exit has a circular section and the combustor entrance has a rectangular one. This work aims to present the process of characterizing the cold flow along the SCTB of the Institute for Advanced Studies (IEAv) using the schlieren technique. The interference of the transition piece in obtaining the required flow conditions at the exit of the SCTB or the entrance of the combustor was mainly evaluated.

Experimental Scramjet Combustion Modes of Hydrocarbon Mixtures at Mach 8 Flight Conditions

Shock tunnel experiments are presented that examine combustion modes of ethylene and a 64/36% ethylene/methane mixture under Mach 8 flight-representative conditions within an axisymmetric scramjet cavity combustor. Combustion modes (scram-mode and dual-mode) are presented via streamwise pressure measurements and visualized via planar laser-induced fluorescence (PLIF) of the hydroxyl (OH) combustion radical in the combustor exhaust plume. Ethylene-fuelled experiments developed scram-mode combustion under lean fuelling, whereas increased fuelling anchored combustion to the fuel injector wake. Further fuelling increases to stoichiometric levels developed dual-mode combustion. Exhaust OH PLIF indicated ethylene scram-mode combustion developed combustion products primarily in the wall-adjacent regions, whereas asymmetric flowfields were noted for the dual-mode. The addition of methane inhibited reactions, with no scram-mode combustion observed at even fuel-rich conditions. Higher fuelling rates developed dual-mode combustion with comparable pressure rises to the ethylene-fuelled dual-mode cases. OH PLIF indicated reduced concentration of OH for the ethylene/methane mixture. This work confirms the retarding influence of methane upon the reaction pathways of ethylene and reaffirms that fuels do not react independently in gaseous mixtures.

Experiments on water vapour condensation within supersonic nozzle flow generated by an impulse tunnel

An impulse facility for analysis of water vapour nozzle flows using both shock tunnel and Ludwieg tube operating modes has been developed and tested at the University of Southern Queensland. Unique high-speed flow visualisation of the water vapour condensation shock has been acquired in the throat region of a nominally two-dimensional convergent-divergent nozzle with a 120 mm2 throat area. This paper presents the facility performance and the time-resolved visualisation results for the nozzle flow, and the results are analysed with the aid of image post-processing tools and quasi-one-dimensional (Q1D) thermodynamic calculations. The experiments have produced qualitative and quantitative data on the flow conditions at four different relative humidity values from 25% to 100%, and over a range of nozzle supply temperatures between 293 K and 343 K. The direct measurement of the pressures and wave speeds within the tunnel resulted in a good agreement with the ideal gas Q1D calculations, particularly for the first incident shocks and expansion waves. Differences progressively increased with the number of reflections, due to the effect of viscous dissipation and the non-ideal end-wall interaction caused by the presence of the nozzle inlet. Moreover, flow visualizations of the location and orientation of both the weak disturbance waves and the condensation shock enabled the investigation of the supersonic water vapour condensation within the nozzle of this impulse facility. Based on these measurements, the nozzle flow downstream of the condensation shock was demonstrated to have a lower Mach number relative to the dry nitrogen flow case. This phenomenon is due to the phase change heat release, which appeared to be more significant when the condensation shock intensity was higher. Although further development of the apparatus is possible, the impulse test-rig configuration that we have developed permits a relatively high flow rate of test gas for a short period of time, leading to significantly reduced operating expenses relative to a continuous flow facility with the same mass flow rate. It is therefore suggested that future studies further improve and test the impulse facility techniques for studying condensation phenomena as a potential alternative to high-power, continuous flow facilities.

Academician Yu Hongru: Pioneer of New Detonation Driving Method for Shock Tubes

Yu Hongru studied mathematics and mechanical engineering in his early years, laying a solid foundation for the combination of science and engineering. In 1956, he was admitted to the Graduate School of Mechanics, Chinese Academy of Sciences. Under the guidance of Qian Xuesen, Lin Tongji, and other masters, as well as his inherent quality and understanding, he embarked on an experimental research journey of "shock wave phenomenon and its application" . Yu Hongru abandoned the techniques used in other countries to meet the needs of our country and explored a low-cost and efficient high-enthalpy experimental method detonation drive, which was widely praised by his international peers. He also actively promoted the application of shock tube and shock tunnel technology in national defense, aerospace engineering, and civil engineering. At the beginning of the 21st century, he researched the performance of a hypersonic ramjet and explored a new concept of using a catalytic composite method to improve the thrust of ramjets. Yu is also actively concerned about the development of China’s Pneumatic Center. He paved the way for the later reseachers.

Effect of the axial cavity with an opposing high-pressure jet combination in a Mach 6 flow condition

Flow control techniques mitigating the aerothermal heating over bodies are essential for long-duration hypersonic flight. The combination technique of a nose cavity with an opposing jet is exhibited promising effects. There is a lacuna of experimental studies on unsteady aspects of the flow and the heat transfer effects within the cavity. In the current study, we investigated the heat transfer variations and nature of shock around a blunt body with a cylindrical and parabolic cavity geometries combined with opposing jet. Experiments are conducted in the hypersonic shock tunnel for a flow Mach number of 6 for an opposing jet pressure ratio of 18. Time-resolved high-speed schlieren imaging and heat flux measurements over the model surface and within the cavity are carried out. The cavity flow is established by verifying the measured heat flux, pressure, and shock oscillation results. The combination technique shows a significant reduction of heat transfer and flow oscillations in comparison to the nose cavity technique. Notably, the parabolic cavity with the opposing jet showed about a 33% reduction of surface heat flux and a 43% reduction in flow oscillations. The substantial increase of heat flux due to the shock interaction is noted on the nose surface for a cylindrical cavity geometry, and such effect was not prevalent in the case of flow with parabolic cavity geometry.

Shock–Boundary Layer-Interaction Control Through Recirculation in a Hypersonic Inlet

Abstract This technical brief presents a flow separation mitigation device, called cavity-recirculator that can be used to control flow separation during shock wave–boundary layer interaction (SBLI) in high-speed intake flows. The cavity-recirculator isolates the flow separation bubble generated at the SBLI spot, thereby thinning the boundary layer and reducing the blockage of the inviscid stream in the duct. The device has a potential application in scramjet engine intakes/isolators. The cavity-recirculator was tested on a single-ramp-compression intake model in a hypersonic shock tunnel, in a freestream of Mach 8 (±0.1). The device operation and effectiveness were assessed by flow visualization and pressure measurements in the test model. The measurements and visualization displayed a mitigation of flow separation through an improved flow field with a single shock train and the absence of flow separation shock, in the inlet.

Nonlinear resonance of low- and high-frequency waves in a cooled hypersonic boundary-layer

We investigated non-linear resonance of low- and high-frequency waves (LHFWs) in a cooled hypersonic boundary layer. First, an experimental campaign was conducted in the high-enthalpy shock tunnel at the Japan Aerospace Exploration Agency using a circular-cone model. The results allowed us to identify early-developing low-frequency waves and strong cross-bicoherence among LHFWs, indicating that non-linear phase couplings existed. Second, to explain the mechanisms involved, stability analyses were performed under the test conditions. In a region upstream of the experimental transition point, a linear analysis of the frequencies of dominant fluctuations in the experiment clarified the existence of several wave triads that satisfy a resonance condition, requiring phase synchronization among three waves. Assuming that the initial amplitude of the low-frequency mode was higher than the initial amplitudes of the high-frequency modes, a weakly non-linear analysis was conducted. The results suggest that as the initial amplitude of the low-frequency mode was increased, the remaining two high-frequency modes deviated from the linear behaviour: the mode at lower frequency was slightly destabilized, whereas the higher-frequency mode was stabilized. It followed that the spectral broadening to lower-frequency bands was relatively enhanced.

A Fast-Response Calorimeter with Dynamic Corrections for Transient Heat Transfer Measurements

Robust fast-response transient calorimeters with novel calorimeter elements have attracted the attention of researchers as new synthetic materials have been developed. This sensor uses diamonds as the calorimeter element, and a platinum film resistance is sputtered on the back to measure the temperature. The surface heat flux is obtained based on the calorimetric principle. The sensor has the advantages of high sensitivity and not being prone to erosion. However, non-ideal conditions, such as heat dissipation from the calorimeter element to the surroundings, can lead to measurement deviation and result in challenges for sensor miniaturization. In this study, a novel transient calorimeter (NTC) with two different sizes was developed using air or epoxy as the back-filling material. Numerical simulations were conducted to explain the complex heat exchange between the calorimeter element and its surroundings, which showed that it deviated from the assumption of an ideal calorimeter sensor. Accordingly, a dynamic correction method was proposed to compensate for the energy loss from the backside of the calorimeter element. The numerical results showed that the dynamic correction method significantly improved the measurement deviation, and the relative error was within 2.3% if the test time was smaller than 12 ms in the simulated cases. Detonation shock tunnel experiments confirmed the results of the dynamic correction method and demonstrated a practical method to obtain the dynamic correction coefficient. The accuracy and feasibility of the dynamic correction method were verified in a single detonation shock tunnel and under shock tube conditions. The NTC calorimeter exhibited good repeatability in all experiments.

Influence of shock attenuation on tailored operation in free piston shock tunnels

The free piston shock tunnel is a type of shock tunnel with high performance. For this type of tunnel, the influence mechanism of shock wave attenuation on tailored operation is explored by numerical simulation and theoretical analysis. By introducing the normalized velocity, the simple constraint equation for shock wave under the tailored operation is deduced. Moreover, the real gas effect is also taken into account in this equation. Based on the equation, the tailored operation of shock tunnels can be predicted with very few calculations. The present study shows that the change rate of the thermodynamic state of the gas behind the shock wave is inconsistent with the attenuation rate of the shock wave, which is the fundamental reason why the wind tunnel achieves tailored operation at a lower Mach number of shock waves. This lower Mach number of shock waves differs from the corresponding ideal value by a factor, which is about the square root of shock attenuation rate.

Impulse force-measurement system

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.

Skin friction estimation on a surface under shock-boundary layer interaction

This article presents correlations for indirect measurement of skin friction inside a laminar separation bubble induced by hypersonic shock-boundary layer interaction (SBLI) on a flat plate. The correlations, based on parameters that are known to influence the SBLI region, were developed using exhaustive numerical and analytical studies. Experiments were conducted in a hypersonic shock tunnel at Mach 8.6 (±0.22) to measure surface heat-flux and pressure in the zone of SBLI on a flat plate, which were then used to supplement and validate the correlations. The data predicted by the correlations agreed reasonably well with that of exact solutions. The case studies contained non-reacting air, behaving as a perfect gas on a flat surface.

Dual/scram-mode combustion limits of ethylene and surrogate endothermically-cracked hydrocarbon fuels at Mach 8 equivalent high-enthalpy conditions

This paper examines the scram/dual-mode combustion limits of hydrocabon fuels within a Mach 8, scramjet combustor. Flight-equivalent flows were delivered to the axisymmetric, cavity combustor via a reflected shock tunnel. Two scramjet fuels were examined: ethylene and a surrogate mixture representing endothermically cracked n-dodecane. Combustion modes were examined via static pressure sensors and through both chemiluminescence imaging, and planar laser induced fluorescence (PLIF) of the OH combustion radical in the combustor exhaust plume. Ethylene-fuelled experiments developed scram-mode combustion under reduced fuelling conditions, experiencing shock wave dominated flowfields. OH PLIF diagnostics indicated such combustion modes developed a ring-like structure of combustion products, primarily axisymmetrically adjacent to the combustor wall. Increased fuelling anchored combustion downstream of the fuel injector, while further increases instigated dual-mode combustion. In this mode, subsonic combustion regions combine with the supersonic coreflow to permit the transfer of information upstream with substantially increased pressure encountered. Optical diagnostics indicate broadly asymmetric, unsteady combustion features. The surrogate mixture representing endothermically cracked n-dodecane experienced rapid onset from no-combustion (optically confirmed) to fully developed dual-mode combustion at critical fuelling rates. OH PLIF signals and chemiluminescence of this fuel were weaker than comparable ethylene cases, indicating potential differences in combustion pathways.

충격파 터널에서의 가속도계 기반 항력 측정

충격파 터널에서 항력 측정을 위하여 가속도계 기반 시스템을 설계 및 구축하고, 마하 6 유동조건에서 원뿔형 모델의 항력계수를 측정하였다. 항력 측정 시스템의 마찰력을 보정하기 위하여 간단하고 직관적인 교정 방식을 제시하고, 항력 측정 결과를 단순 원뿔형 모델의 전산유체역학 해석 결과와 비교, 분석하였다. 다양한 형상의 지지대를 이용하여 지지대에 의한 항력 측정 간섭 영향을 파악하고 이를 최소화하기 위한 디자인을 제시하였다. 수정된 지지대를 이용하여 항력계수를 측정한 결과, 지지대에 의한 항력계수 오차가 작아지는 것이 확인되었다.