State Key Laboratory of High Temperature Gas Dynamics

The effect of wall temperature on the size of the separation bubble in the shock wave/turbulent boundary-layer interaction of a 24 deg compression ramp with Mach 2.9 is numerically investigated. The ratios of wall temperature to recovery temperature are 0.6, 1.14, 1.4, and 2.0, respectively. To validate the simulation, the statistical results with are tested and the results show a good agreement with theoretical and experimental results. It is shown that wall temperature has a remarkable effect on the size of the separation bubble and the size increases significantly with the increase of wall temperature. Through theoretical analysis, combined with numerical results, we get a semitheoretical formula , in which and are the length of the separation bubble and the thickness of upstream boundary layer, respectively. The turbulent kinetic energy budgets are also analyzed based on the numerical data, and results show that turbulence kinetic energy is chiefly produced both in the buffer layer and near the shock wave, and turbulent dissipation is mainly in the center of the separation bubble as well as in the near-wall region. It is also shown that the intrinsic compressibility effect is not significant in all these cases.

Experiments were conducted to characterize shock train oscillation under the simultaneous variation of the incoming Mach number and backpressure. Under steady and low-frequency oscillatory backpressure (2 Hz), the incoming Mach number varied from 1.8 to 2.4. According to the intersection of downgoing background wave with bottom front leg, Mach stem, and top front leg of the normal shock train leading edge, the normal shock train/background wave interaction can be divided into three types. Two types of oblique shock train/background wave interaction exist. The downgoing (upgoing) background wave upstream of the oblique shock train can cause the upgoing (downgoing) shock in the shock train leading edge to become the dominated shock. Two modes of shock train oscillation were found: oscillation mode 1, in which the shock train oscillated in the favorable gradient region of the relaxing boundary layer, and oscillation mode 2, where the shock train enters the adverse pressure gradient region caused by the impingement of background wave. Compared with mode 1, mode 2 leads to a larger upstream movement of the shock train and more intense pressure fluctuation. The oscillation of the shock train is caused by instability in the separation region behind the shock train leading edge. The oscillatory backpressure only affected the motion of shock train during each oscillation period. The overall movement trend of shock train is determined by the incoming Mach number and the mean value of backpressure. The increase of incoming Mach number and backpressure can lead to the enhancement of shock train oscillation.

A system for three-dimensional computed tomography of chemiluminescence was developed to measure flames over a large field angle. Nine gradient-index rods, with a 9 × 1 endoscope and only one camera are used. Its large field of view, simplicity, and low cost make it attractive for inner flow field diagnostics. To study the bokeh effect caused by the imaging system on reconstruction solutions, fluorescent beads were used to determine the blurring function. Experiments using a steady diffusion flame were conducted to validate the system. Three models, namely the clear-imaging, out-of-focus imaging, and deconvolution models, were utilized. Taking the bokeh effect into account, the results suggest that based on run-times the deconvolution model provides the best reconstruction accuracy without increasing computational time.

Force tests were conducted at the long-duration-test shock tunnel JF12, which has been designed and built in the Institute of Mechanics, Chinese Academy of Sciences. The performance tests demonstrated that this facility is capable of reproducing a flow of dry air at Mach numbers from 5 to 9 at more than 100 ms test duration. Therefore, the traditional internal strain-gauge balance was considered for the force tests use in this large impulse facility. However, when the force tests are conducted in a shock tunnel, the inertial forces lead to low-frequency vibrations of the test model and its motion cannot be addressed through digital filtering because a sufficient number of cycles cannot be found during a shock tunnel run. The post-processing of the balance signal thus becomes extremely difficult when an averaging method is employed. Therefore, the force measurement encounters many problems in an impulse facility, particularly for large and heavy models. The objective of the present study is to develop pulse-type sting balance by using a strain-gauge sensor that can be applied in the force measurement of 100 ms test time, especially for the force test of the large-scale model. Different structures of the S-series (i.e., sting shaped balances) strain-gauge balance are proposed and designed, and the measuring elements are further optimized to overcome the difficulties encountered during the measurement of aerodynamic force in a shock tunnel. In addition, the force tests were conducted using two large-scale test models in JF12 and the S-series strain-gauge balances show good performance in the force measurements during the 100 ms test time.

Abstract The current work presents the collision integral data for N( 4 S )–N( 4 S , 2 D , 2 P ) and O( 3 P , 1 D , 1 S )–O( 3 P , 1 D , 1 S ) interactions in the temperature range of 500–50 000 K. The collision integrals are calculated based on high-quality potential energy curves (PECs) obtained from fitting the high-level <mml:math xmlns:mml="http://www.w3.org/1998/Math/MathML" overflow="scroll"> <mml:mi>a</mml:mi> <mml:mi>b</mml:mi> <mml:mtext> </mml:mtext> <mml:mi>i</mml:mi> <mml:mi>n</mml:mi> <mml:mi>i</mml:mi> <mml:mi>t</mml:mi> <mml:mi>i</mml:mi> <mml:mi>o</mml:mi> </mml:math> calculation data in a wide energy range to the neural network (NN) functions. In the construction of PECs, the diabatic PECs are adopted when avoided crossings exist because the diabatic paths are much more likely to be followed for such situations. Moreover, the nonadiabatic transition effects are estimated to be negligible for PECs crossings. The accuracy of traditional analytical formulas to fit PECs are also examined. It is found that the collision integral calculations are sensitive to the accuracy of PECs and the NN based PECs overwhelm the others. The contribution of inelastic excitation exchange processes to the diffusion collision integrals are also computed by using an accurate evaluation of the differences of PECs for gerade and ungerade pairs of excited atoms. Finally, based on the new collision integral data, we calibrate the collision model parameters suitable for the widely used particle simulation methods. The collision integrals and collision models developed in this work can be used to support high-confidence simulations of weakly ionized air plasma problems.

The strong coupling interactions of non-equilibrium flow, microscopic particle collisions and radiative transitions within the shock layer of hypersonic atmospheric re-entry vehicles makes accurate prediction of the aerothermodynamics challenging. Therefore, in this study a self-consistent non-equilibrium flow, collisional–radiative reactions and radiative transfer fully coupled model are established to study the non-equilibrium characteristics of the flow field and radiation of vehicle atmospheric re-entry. The comparison of the present calculation results with flight data of FIRE II and previous results in the literature shows a reasonable agreement. The thermal, chemical and excited energy level non-equilibrium phenomena are obtained and analysed for the different FIRE II trajectory points, which form the critical basis for studying the heat transfer and radiation. The non-equilibrium distribution of excited energy levels significantly exists in the post-shock and near-wall regions due to the rapid vibrational dissociation and electronic under-excitation, as well as the wall catalytic reactions. The analysis of stagnation-point heating of FIRE II illustrates that the translational–rotational convection and the dissociation component diffusion play key roles in the aerodynamic heating of the wall region. The spectrally resolved radiative intensity in the entire flow field indicates that the vacuum ultraviolet radiation caused by the high-energy nitrogen atomic spectral lines makes the main contribution to the radiative transfer. Finally, it is found that the non-equilibrium flow–radiation coupling effect can exacerbate the excited energy level non-equilibrium, and further affect the gas radiative properties and radiative transfer. This fully coupled study provides an effective method for reasonable prediction of atmospheric re-entry flow and radiation fields.

Multiphase Richtmyer–Meshkov instability (RMI) is often accompanied by a dispersed phase of particles, where the evolution of the mix zone width (MZW) is a significant issue. The Stokes number (St) is a key dimensionless parameter for particle-containing multiphase flows because it represents the ability of particles to follow the fluid. However, our theoretical analysis and numerical simulation indicate that the Stokes number is not the only dominant parameter for the evolution of multiphase RMI. This study uses the derivation of particle and fluid momentum equations to demonstrate the inability of the Stokes number to predict MZW evolution, that is, even at the same Stokes number, increasing the particle density or the radius leads to completely different MZW evolution trends. This study proposes a novel dimensionless number, Sd, to measure the effect of drag on the fluid owing to the particles. Sd is the ratio of the relaxation time of the fluid velocity affected by the particle force to the characteristic time of the shock wave. We developed theoretical models of MZW at different Sd values. Subsequently, a set of multiphase RMI numerical simulations on uniformly distributed particles with different St and Sd values was conducted. The numerical results verify the theoretical predictions and effectiveness of the proposed dimensionless number. The phase diagram containing different simulation cases demonstrates that the Stokes number cannot be used to predict MZW and must be combined with Sd to determine its evolution.

An experimental investigation of coupling small strut and cavity for enhancing supersonic mixing and combustion had been carried out on direct-connect supersonic combustion test facility. The wall static pressure, CH* chemiluminescence imaging and the schlieren system were used to research the effect of different strut type and jet. For long distance of strut to cavity, when the fuel injection was in a supersonic flow , the effects of combustion enhancement was obvious and was weakened on subsonic mode. For short distance of strut to cavity, combustion enhancement effect is very obvious, significantly increase the volume of high-temperature flame near the cavity.

Abstract Internal temperature monitoring of high-speed propulsion systems is highly important for engine performance evaluation and lifetime prediction. As a passive optical measurement method without the need for an external light source and without flow field interference, the emission spectrum measurement technique has good application prospects for harsh measurement environments. As the main combustion product, high-temperature water vapor shows a strong emission intensity that is highly suitable for temperature measurement applications. We propose use of the band integral ratio to remove the high resolution measurement requirements for the spectrum acquisition system. In addition, the temperatures of methane-oxygen flames with different equivalent ratios are measured successfully under the condition that the influence of self-absorption on the measurements is considered.

Combustion has been widely studied in the literature, but very little work was focused on the microscopic level. In this paper, the DSMC method is applied to simulate the microscopic behavior of the spontaneous combustion of hydrogen oxygen mixture. It is found that the ignition delay time of the mixture depends on many factors, such as the physical size, temperature, pressure, and dilution. Comparison between DSMC and CFD results shows that more atomic hydrogen is consumed through reaction HO2+H→H2+O2 at temperature close to the extended second explosion limit due to localized distribution of reactants, which may indicate the importance of microscopic behavior on low temperature combustion.

The flame dynamics in the combustion chamber of a hybrid rocket motor were visualized using novel chemiluminescence imaging. A multidirectional visualization system employing [Formula: see text] endoscopes generated images based on methylidyne chemiluminescence (CH*), with one endoscope in the precombustion chamber and two in the postcombustion chamber. Images were collected with a high-speed camera using a 1 ms exposure and a 1 kHz frame rate. Fuel grains having a helical or a conventional circular port structures were assessed, and combustion trials were conducted using a laboratory-scale hybrid rocket motor with oxygen as the oxidizer at mass flow rates from 10.43 to [Formula: see text]: equivalent to combustion chamber pressures ranging from 0.7 to 1.24 MPa. Flame structures were observed during the ignition, combustion, and shutdown stages; and the helical grain generated a larger, more intense flame zone. A proper orthogonal decomposition analysis showed that the helical grains also produced a greater degree of turbulence and stronger oscillations. These results confirm that a helical structure increases the flow turbulence and convective heat transfer in the combustion chamber. These effects lead to higher regression rates and better mixing efficiency that may, in turn, provide greater combustion efficiency at optimized oxidizer/fuel ratios.

Aluminum and boron nanoparticles are added to kerosene in this paper to improve the combustion properties of kerosene. When the particle concentration is 30 g/L, the volumetric heat value of fuel increases by 3 and 5%, respectively, with the addition of aluminum and boron nanoparticles. Meanwhile, combustion experiments in a supersonic combustor are conducted to study the combustion characteristics of Al-kerosene nanofuels and B-kerosene nanofuels. The air flow rates of all experiments are about 1.77 kg/s, the total temperature is 1500 K, and the total pressure is 1.2 MPa. The combustion flow and flame structures of kerosene with different particles and different concentrations are studied, and the results indicate that the addition of nanoparticles to the fuel enhances combustion and heat release enhanced, and the flame stabilization mode is changed from the cavity stabilization to the jet-wake stabilization mode with a higher particle concentration. Meanwhile, the unsteady characteristics of flame are studied. The flame oscillation is intensified, and the fundamental frequency of the flame increases with the increase in particle concentration. Besides, the addition of nanoparticles significantly improves the combustion efficiency of kerosene. When the particle concentration is 30 g/L, the combustion efficiencies of nanofuels are increased by 15% and 17.5%, respectively, with the addition of aluminum and boron nanoparticles.

Abstract The choice of electrode configuration and dielectric material is critical to the discharge process and plasma characteristics of a dielectric barrier discharge (DBD) reactor. In this study, a new electrode configuration of DBD reactor with copper mesh as electrode inserted between dielectrics is proposed, which has a much higher capacitance than the conventional double-dielectric layer DBD reactor. Two materials with different relative dielectric permittivities, alumina and zirconia, are chosen as dielectrics for an experimental comparison of CO 2 decomposition. The experimental results show that the conversion rate of CO 2 for the reactor with copper mesh inserted between dielectrics are higher than that of the corresponding double dielectric layer reactor under the same discharge power, and the conversion rate of CO 2 with zirconia as a dielectric material is higher than the case of alumina as a dielectric. Further analysis of discharge characteristics shows that for the reactor with copper mesh inserted between dielectrics, the applied voltage required for discharge is significantly reduced, the amount of transferred charge is significantly increased, and the number of micro-discharge current pulses as well as the average lifetime during a single voltage cycle are also considerably increased, leading to an increase in the CO 2 discharge efficiency and conversion rate.

To develop a more advanced 3D computed tomography of the chemiluminescence method, the first quantitative 3D diagnosis was realized. The nonlinearity coefficient, the nonuniformity coefficient of the camera response, and various optical fiber attenuation coefficients were obtained through correction experiments. The conversion relationship between the number of photons released by the target object per unit time and the camera gray value at a specified solid angle was also calibrated. To verify the quantitative reconstruction equation, 3D reconstructions of a methane-air flat flame and a simulated phantom were performed for comparison. The method can overcome artificial distortions caused by uncorrected reconstruction.

The need to increase the payload capacity of the rockets motivates the development of high-power rocket engines. For a chemical propulsion system, this results in an increasing thermal load on the structure, especially the combustion chamber and nozzle must be able to withstand the extreme thermal load caused by high-temperature and high-pressure combustion gas. In order to protect the structure from the effect of increasing heat flux, it is necessary to counteract such effect with more advanced thermal management technology. This requires us to accurately predict the aerodynamic heating of the structure by high-temperature and high-speed combustion gas. In this study, a high-temperature combustion gas tunnel developed in the laboratory is used to produce high-speed combustion gas. Combined with the results of numerical calculation, the flow and aerodynamic heating characteristics of air and hydrogen–oxygen combustion gas under the same total temperature and pressure are analyzed and compared. The comparison revealed that the combustion gas flow in the nozzle has higher static temperature, velocity, and smaller Mach number. When the combustion gas flows around the sphere, the shock standoff distance and stagnation pressure are smaller than those of air, and the wall heat flux is much larger than that of air. The active chemical reaction in the combustion gas makes the aerodynamic heating of the structure more severe. Finally, through the analysis of a large amount of data, a semi-empirical formula for the heat flux of the stagnation point heated by a high-speed hydrogen and oxygen equivalent ratio combustion gas is obtained.

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.

A dynamic-force extraction, based on the least-squares method, is proposed for micro-propulsion testing. Having modeled the displacement oscillation of a micro-newton torsional pendulum, the time evolution of the dynamic force may be calculated if the stand constants are well calibrated. According to the linear characteristic of the motion equation, a reconstruction of the dynamic thrust reduces to solving linear equations. The simulation analysis shows that the error is affected by the sensor noise and the low-pass filter as well as the sampling rate. Validation experiments were performed showing that this method reconstructs the dynamic force well up to 8 Hz with an error less than 15 μN. The noise-induced error moreover varies little with frequency.

To develop ammonia-fueled micro gas turbines, an industrial-grade full-size annular combustion chamber and a gas injector specifically designed for NH 3 combustion were manufactured using 3D printing. Precracking NH 3 is proposed as an ideal solution to overcome the fuel’s low flammability. This study experimentally and numerically investigated the combustion characteristics of precracked NH 3 in the annular combustion chamber. The combustion zone maintained an equivalence ratio of 0.54 with a 0.5 MPa pressure. Experimental results revealed that even at 838 K air temperature, 20 wt % precracked NH 3 failed to sustain stable combustion. However, stable combustion was achieved when precracked degree reached 30 wt % with 560 K air and 40 wt % with ambient-temperature air. To match CH 4 ’s ignition delay time and laminar burning velocity, the required NH 3 precracked degrees were approximately 7 and 35%, respectively. The tests achieved 94.5% combustion efficiency with minimal NOx emissions of 376 ppm, though emissions increased sharply with rising initial air temperatures. Numerical analysis indicates that NO generation strongly correlates with N and HNO radicals, while initial air temperature exerts greater influence than the degree of NH 3 precracked on NO formation. The equivalence ratio in the pilot zone exerts a profound influence on the NO emission concentrations. These findings provide valuable guidance for the stable operation of NH 3 -fueled gas turbines, thereby facilitating relevant research.

A technique for optical emission spectroscopy is developed and deployed in the detection of cooling-water leakage in an arc-heated wind tunnel at China Academy of Aerospace Aerodynamics. The arc heater operates at temperatures of 5000–9000 K and pressures of 2–6 atm. Two mass-average enthalpy conditions, (test 1) and (test 2), are studied. From the spectral characteristics of the normal and leakage spectra, the 656.28-nm emission spectral line of atomic hydrogen and the 777.19-nm emission spectral line of atomic oxygen are selected for coolant-leakage detection in the arc heater. Given the emission intensity ratio of and , detection limits for the mass flow of the leaking water are derived; for tests 1 and 2, they are and , corresponding to equilibrium temperatures ranging from 6000 to 8000 K and 5500 to 7500 K, respectively. This work demonstrates the feasibility and potential of the optical emission spectroscopy technology in high-enthalpy arc heater health diagnosis, especially in regard to the coolant leakage diagnosis.

This study investigates the flowfield patterns and distributions of surface heat flux due to interactions among three-dimensional shock waves at the junction of the body and wing of an aircraft by solving Reynolds–averaged Navier–Stokes equations at a Mach number of 10 and attack angles ranging from 5° to 20°. The results indicate that the structures of wing/body-shock interactions vary significantly with test conditions. Four types of shock interaction patterns were observed: interaction-free, type I regular, type II regular, and Mach interactions. Once the flowfield of the shock interactions had been established, aerodynamic heating loads of the wing and body were affected by the flowfield structures. Wing/body-shock interactions produced uneven heat flux distributions on the surface and caused an abnormally high heat flux at a localized position. Five profiles of the distribution of heat flux were extracted to describe its characteristics on the surface according to the position and magnitude of the peaks of the localized heat flux. Induction-related factors that led to the peaks were classified into three types: reflected shock/boundary-layer interaction, contact surface impinging, and contact surface grazing.