Mach Number Increases Research Articles

Comparison of Dual-Combustion Ramjet and Scramjet Performances Considering Combustion Efficiency

The performances of a dual-combustion ramjet (DCR) and a scramjet were compared via computational fluid dynamics numerical simulation to provide theoretical guidance for engine selection for a hypersonic vehicle. Kerosene, C12H23, with an equivalence ratio of 0.8, was employed as the fuel, and the reactive flow was modeled using six-species and four-step chemistry. The results show that the DCR has a central combustion mode, which has a smaller temperature gradient and more uniform heat release, resulting in higher combustion efficiency, compared to the near-wall combustion mode of the scramjet. The total pressure recovery coefficient of scramjet is 0.9% lower than that of DCR under the Ma6 condition, but 5.6% higher than that of DCR under the Ma7 condition. The combustion efficiency of DCR is 35.6% and 25.4% higher than that of the scramjet under Ma6 and Ma7 conditions, respectively. The decrease in the combustion efficiency of the DCR is caused by the increase in the dissociation rate of CO2 into CO with the increase in temperature. The performance of DCR is better than that of scramjet under both conditions. However, the performance advantage of DCR decreases as the Mach number increases. Specifically, under the conditions of Ma6 and Ma7, the specific impulse or specific thrust of DCR was 2.67 times and 1.51 times that of scramjet, respectively.

Analysis of stagger effects on Busemann supersonic biplane airfoil in shock tube tests by point diffraction interferometer method

The Busemann biplane concept is proposed for low-boom low-drag supersonic aircraft. Studies on the supersonic biplane have been focused on supersonic flow, whereas the understanding of transonic aerodynamic characteristics is essential for aircraft design. Significantly, the complicated flow field with a shock and boundary layer interaction between Busemann biplane elements in transonic flow remains unclear. In the present study, the transonic aerodynamic characteristics of the two-dimensional Busemann biplane are investigated by shock tube tests and numerical simulations, focusing on the flow field between the biplane elements. In addition, the stagger effects, which change the relative position of the lower and upper elements in the freestream direction to avoid the choked flow between the biplane elements, are also analyzed. A shock tube system creates a test flow at Mach numbers of 0.6 to 0.8, and the Reynolds number is set at 2.85 × 105. The Point Diffraction Interferometer (PDI) method is applied to determine the density distribution around the model and the pressure coefficient on the surfaces of the biplane elements. Symmetrical density distribution is confirmed in the Busemann biplane (the baseline model, or the biplane with no stagger). A normal shock wave is generated between the biplane elements. When the Mach number increases, the normal shock wave moves downstream. The normal shock position is quantitatively stable with small fluctuation amplitudes. The fringes are strongly curved near the surfaces of the biplane elements due to the flow separation near the airfoil surfaces. When the Mach number increases, the separation area increases. When the Mach number increases, the peak value of the pressure coefficient on the surfaces of the biplane elements increases. On the other hand, the normal shock wave between the biplane elements is eliminated in the staggered models. The changes in density distribution between the biplane elements are minor compared to the case of the baseline model, which indicates that the stagger reduces the choked flow between the biplane elements. The peak value of the pressure coefficient on the wing surface increases when the stagger value increases. The stagger effects on reducing the total drag coefficient of the biplane models are confirmed.

Effect of the flame motion on azimuthal combustion instabilities

In the analysis of combustion instabilities in the annular combustors, especially for the azimuthal combustion modes, the flame motion is usually neglected and the flame is typically regarded as a discontinuity at rest. However, this simplicity may cause severe prediction errors for certain conditions. In this paper, the influence of the flame motion on the entropy, vortex, acoustic, and intrinsic modes in the annular combustion chamber is investigated by constructing an analytical model with the flame motion. Meantime, the influence of flame model parameters is also explored. The flame motion can affect both the frequency and growth rate of the entropy mode, but the growth rate is influenced more significantly. This effect can be damped or enlarged by the inlet Mach number depending on which parameter of flame model remains unchanged. The effect of the flame motion on vortex, acoustic, and intrinsic modes remains small when the inlet Mach number is very small. However, the increase of inlet Mach number can enlarge this effect.

Interactions between upstream-propagating guided jet waves and shear-layer instability waves near the nozzle of subsonic and nearly ideally expanded supersonic free jets with laminar boundary layers

The interactions between upstream-propagating guided jet waves and shear-layer instability waves near the nozzle of subsonic and nearly ideally expanded supersonic, isothermal free jets are investigated for jets at Mach numbers between 0.50 and 2 with fully laminar exit boundary layers of different thicknesses. The velocity spectra in the shear layers downstream of the nozzle exhibit strong narrow peaks for the first azimuthal modes, associated with growing Kelvin–Helmholtz instability waves. The frequencies of the predominant peaks are close, but not necessarily equal, to those of the most amplified instability waves predicted from the mean flow fields using linear stability analysis. They also fall in most cases within or very near the allowable frequency bands of the free-stream upstream-propagating guided jet waves obtained using a vortex-sheet model and jump from one band to another as the Mach number increases. At these frequencies, moreover, high levels organized into elongated stripes are found in the jet potential core and standing-wave patterns are visible at the edges of the shear layers in the power spectral densities of pressure and velocity fluctuations. Therefore, the free-stream upstream-propagating guided jet waves appear to interact with and excite the instability waves near the nozzle of the present jets, as in screeching and impinging jets. This explains the disparities of the frequencies and azimuthal modes of the instability waves dominating early on in the shear layers of initially laminar jets and their discontinuous changes and staging behaviours as the jet velocity varies.

Compressibility effects on statistics and coherent structures of compressible turbulent mixing layers

The effects of compressibility on the statistics and coherent structures of a temporally developing mixing layer are studied using numerical simulations at convective Mach numbers ranging from $M_c=0.2$ to $1.8$ and at Taylor Reynolds numbers up to 290. As the convective Mach number increases, the streamwise dissipation becomes more effective to suppress the turbulent kinetic energy. At $M_c=1.8$ , the streamwise dissipation increases much faster than the other two components in the transition region, even larger than pressure–strain redistribution, correlating with the streamwise elongated vortical structures at a higher level of compressibility. We confirm the existence of the large-scale high- and low-speed structures in the mixing layers, which accompany the spanwise Kelvin–Helmholtz rollers at low convective Mach number and dominate the mixing layer at higher convective Mach number. Conditional statistics demonstrate that the large-scale low-speed structures are lifted upwards by a pair of counter-rotating quasi-streamwise rollers flanking the structures. The small-scale vortical structures have an apparent preference for clustering into the top of the low-speed regions, which is directly associated with high-shearing motions on top of the low-speed structures. The high-speed structures statistically exhibit central symmetry with the low-speed structures. The statistics and dynamics of large-scale high- and low-speed structures in the compressible mixing layers resemble those in the outer region of the turbulent boundary layers, which reveals the universality of the large-scale structures in free shear and wall-bounded turbulence. A conceptual model is introduced for the large-scale high- and low-speed structures in turbulent mixing layers.

Energy and performance analysis of a turbofan engine with the aid of dynamic component efficiencies

Assessment of performance of turbofan engine with different design parameters is crucial for meeting the performance requirements by considering each key components of the engine. To comprehend influences of component efficiencies to mitigate environmental effect from turbofans has been hot topics in aviation field recently. In this study, firstly, impacts of polytropic efficiencies of fan, compressor and turbine as well as pressure ratio of combustor (CPR) on several turbofan performance are dealt with at several flight conditions. Secondly, it is tried to show difference of performance metrics under ideal and real conditions. The discrepancy between performance parameters computed at ideal and real cases gets relatively high. Namely, at take-off condition, the difference between ideal and real specific fuel consumption is computed as 29.12% whereas it is found as 28.37% at cruise condition, which shows that considering the system as ideal makes the computations inappropriate for performance analysis. Moreover, performance parameters of turbofan is more sensitive to compressor efficiency compared with turbine. As the polytropic efficiencies of fan and compressor are close to highest, net thrust of the engine develops from 109.05 kN (baseline) to 124.71 kN at take-off while it increases from 26.36 kN (baseline) to 29.27 kN at cruise condition. With effect of the elevated pressure ratio of combustor and efficiency of turbine, thrust of the engine increases to 118.23 kN at take off and to 27.84 kN at cruise condition. Finally, as Mach number increases, the difference between ideal and real performance values sharply increases. Therefore, when analyzing on turbofan engines, the assumptions should be minimum as possible as, otherwise the findings make the engineers to misguide for system optimization. Besides, these outcomes show that if the components with higher polytropic efficiency can be obtained, overall efficiency of turbofan, thereby environmental sustainability could be elevated to upper level compared with baseline.

Mechanism of shock-train/boundary-layer interaction in spanwise concave isolator channels

A series of spanwise concave channel models of isolators with different radii were investigated to reveal the mechanism of shock train/boundary layer interaction in the inward-turning inlet. Numerical simulations of the shock train flow under different incoming Mach numbers, incoming static pressures, and backpressure ratios were performed. The results show that the shock train leading edge exhibits a “λ” shape in the concave channel, and that a Mach stem is present at the intersection of the shock waves on the top and bottom walls. As the radius of the concave wall is increased, the streamwise force generated by the wall decreases and the upward force perpendicular to the wall increases, which leads to an increase in the thickness of the boundary layer and weakening of the ability of the boundary layer fluid to resist the pressure gradient. Owing to the increase in the radius of the concave channel, the shock train moves upstream, the Mach stem in the shock train leading edge becomes shorter, the distance between the reflection shock waves increases, and the shock train becomes longer. Compared with the concave bottom wall, the separation flow at the flat top wall extends further streamwise. As the incoming Mach number increases, the inertial effect is strengthened, and the shock train evolves towards a normal shock wave structure. As the incoming static pressure decreases, the viscous effect is enhanced, and the flow field of the shock train tends to evolve to reflect more shock waves with compression over longer streamwise distances.

Research on flow field characteristics of nozzle plume for a solid rocket motor

In order to explore the flow field characteristics of nozzle plume for a small solid rocket motor (SRM), the two-dimensional axisymmetric N-S governing equations with component transports and chemical reaction terms are established. Then, the finite volume method of AUSM (Advection Upstream Splitting Method) spatial discretization scheme is applied for numerical discretization, and its solution is solved by time iteration to make the plume flow field steady. This technique is applied to research the flow field characteristics and characteristic parameter distributions of nozzle plume under different inflow conditions. The results are shown that: when considering the reburning reactions of nozzle plume, the temperature of the gas jet core area is much higher than the flow field temperature that the reburning phenomenon is not considered. In addition, the reburning phenomenon mainly occurs in the air-gas mixing and gas development areas, but in the initial core area of gas jet, this phenomenon is not obvious. What’s more, with the increase of altitude, the influence area of the gas jet gradually becomes larger. However, as the incoming flow Mach number increases, the intensity of the change of gas temperature along the nozzle axis weakens, and also the number of wave nodes and the influence region of gas jet gradually decrease.

Modeling Swirl Evolution Inside Annular Nozzles Under Compressible Effect

Swirling flows are widely used to increase fuel mixing in air propulsion systems, and the swirl intensity affects combustion efficiency significantly. We propose a novel one-dimensional model to predict the swirl number and directional Mach numbers for compressible swirling flows inside annular nozzles. Limitations of previous models are addressed via asymmetric wall shear stress, compressible friction factor for boundary-layer thickening, turbulent viscosity, wall heat transfer model, and so on. The proposed model is validated for turbulent flows inside convergent–divergent (CD) and straight annular nozzles. Compared to inviscid modeling, significantly improved predictions are observed. While the effects of compressible friction factor and wall heat transfer model are significant, the effect of turbulent viscosity decreases as inlet Mach number increases. The present model is applied to parametric studies to examine physical effects. For CD annular nozzles, the viscous effects on discharge coefficient and thrust efficiency are examined, and consistent results with previous studies are observed. For straight annular nozzles, the effects of primary parameters are examined. In both subsonic and supersonic cases, the axial changes of azimuthal Mach number depends mainly on the wall shear stress. In a subsonic case, the swirl number decreases to the downstream, and an optimal ratio of inner and outer radii () minimizing the swirl decay rate exists. In a supersonic case, however, the swirl number keeps increasing, and the slope increases monotonically over Gamma.

Narrow-band infrared radiation characteristics of rocket exhaust plume by using correction function related to thermodynamic state

Modeling of plume infrared radiation requires the accurate calculation of gas radiation properties of the plume non-uniform flow field. Aiming at the problem that the Curtis-Godson approximation (CGA) method can give the exact result in the limits of weak lines and strong lines, but overestimates the absorption of the optical intermediates path. This study develops an accurate model to study the infrared radiation of rocket plume. This model calculates narrow-band radiation properties using an equivalent half-width corrected by a correction function, which relates to the thermodynamic state of the no-uniform plume and is followed by radiative transfer calculations. Several Lorentz single lines were adopted to evaluate this correction function by the equivalent width of a fictitious non-uniform gas column, results showed high accuracy, especially at optical intermediates. Simulation of plume reaction flow was implemented by a refined 11-species, 17-reaction chemistry scheme, which can describe the afterburning between the multi-component fuel injected from the nozzle and ambient air. Then, adopting the temperature, pressure and radiating species distribution of the plume flow field, the model can predict accurate spectra that fit the experimental spectrometer data well. Finally, the effects of propellant formulation and flight conditions on plume infrared radiation were investigated: (a) hydroxyl-terminated polybutadiene (HTPB) and nitrate ester plasticized polyether (NEPE); (b) flight altitude; (c) flight Mach number. Results show that the composition of the propellant has a significant effect on the plume afterburning, NEPE plume afterburning effect is stronger, and the total radiation intensity is 61.5% higher than HTPB in 1.5–6.0 μm. The increase of flight altitude weakens the radiation, while due to the expansion of the radiant area, the spectral radiation intensity increases significantly. With the increase of Mach number, the band intensity decreases exponentially, and the plume structure gradually develops into a slender shape, the radiation decreases significantly. The research has significance for the application of target early warning, detection and identification.

The effect of mean flow on the intrinsic thermoacoustic instabilities in the duct and annular combustion chambers

Thermoacoustic instability in the combustion chamber of gas turbines, aero and rocket engines has become a topic concern. In this paper, a new instability phenomenon in the field of the thermoacoustic instability - Intrinsic ThermoAcoustic instabilities (ITA) is studied. The classic thermoacoustic instability is caused by the coupling among acoustic disturbances, flow fluctuations and heat release pulsations. When the influence of acoustic disturbance is removed, there is still a thermoacoustic instability phenomenon in the combustion chamber, which is known as ITA. In the previous research on ITA, the scholars found that the ITA is mainly determined by the flame response, but many scholars neglected the influence of the mean flow and entropy waves. In this paper, the influence of the mean flow on the ITA mode in the duct and annular combustion chambers is studied by constructing a low-order acoustic network model. At the same time, it also explores the influence of other parameters on the ITA mode. The mean flow greatly affects the growth rate of the ITA mode, and has little effect on its frequency in the duct combustion chamber model. This is also the main reason for neglecting the effect of mean flow on the ITA mode in previous studies. In addition, the influence of other parameters on the ITA mode is also reflected in the growth rate. In the annular combustion chamber, a new ITA mode, which does not meet the −π criterion, is found. And the increase of Mach number makes all ITA modes more unstable. The circumferential mode number has a great influence on the high frequency ITA mode. In addition, the ITA mode has a significant effect on the thermoacoustic mode of the system in the annular combustion chamber.

Numerical Simulation of Orbit Controlled Hot-gas Lateral Jet of the triangular plate with rudder

Hypersonic interception weapons have high requirements for the efficiency and accuracy of the control system. As an effective control approach, the thermochemical non-equilibrium effect of jet flow must be considered in order to accurately predict the aerodynamic characteristic. At present, relevant researches are mostly focused on the low-flow rate hot jet, while numerical simulation and wind tunnel test studies on high-flow rate hot jet are rare. In this paper, the interaction flow field of the rail-controlled jet at the altitude of 60Km and 70Km has been simulated by solving the three-dimensional viscous fluid governing equations. The influence of the thermalchemical non-equilibrium effect of the jet interaction on the flow field characteristics and aerodynamic characteristics of the aircraft was analyzed. The primary results show that: (1) The flow fields of the cold-gas jet, frozen-gas jet and hot-gas jet have significant differences under the high-flow rate jet condition, which is mainly reflected the size and the pressure value of the separation region before the nozzle. The separation region length of the cold-gas jet is approximately twice that of the hot-gas jet and the length of the frozen-gas jet is between them; (2)Jet interaction will increase the surface pressure of the aircraft and reduce the skin-friction. When Ma=25, the aerodynamic interaction of hot jet is only about 30% of that of cold jet; (3) Jet interaction on aerodynamic characteristics increases with altitudes. With the increase of Mach number, the difference between the predicted results of hot jet and cold jet is enlarged because of the enhancement of the thermochemical non-equilibrium effect.

Numerical investigations on the deformation and breakup of an n-decane droplet induced by a shock wave

This paper adopts the coupled level-set and volume-of-fluid and the large eddy simulation methods to simulate the deformation and breakup of an n-decane droplet under the action of a shock wave. We aim to investigate the effects of the shock Mach number and droplet diameter on temporary deformation and breakup characteristics at high Weber numbers from 5813 to 22 380. Additionally, special attention is paid to subsequent sub-droplet size distributions, which many researchers generally ignore. The results indicate that the evolution of droplet deformation and breakup in the shear breakup regime generally agrees with the obtained experimental data. Based on the present methods, the physical mechanisms for variations of multiple recirculation zones and the development of Kelvin–Helmholtz instability in wave formation are discussed. Larger shock Mach number and smaller droplet diameter can significantly increase the cross-stream and stream-wise deformations. Moreover, both relaxation and breakup times are directly proportional to the initial droplet diameters but inversely proportional to the shock Mach numbers. Eventually, as the shock Mach number increases, the superficial area and mass ratios of sub-droplets to parent droplets all increase from 5.596 to 8.278 and from 23.38% to 38.38%, while the ratios increase from 2.652 to 18.523 and from 4.63% to 92.7%, respectively, as the droplet diameter decreases.

Implications of weak rippling of the shock ramp on the pattern of the electromagnetic field and ion distributions

Collisionless shocks undergo structural changes with the increase of Mach number. Observations and numerical simulations indicate development of time-dependent rippling. It is not known at present what causes the rippling. However, effects of such rippling on the field pattern and ion motion and distributions can be studied without precise knowledge of the causes and detailed shape. It is shown that deviations of the normal component of the magnetic field from the constant value indicate certain spatial dependence of the rippling. Deviations of the motional electric field from the constant value indicate time dependence. It is argued that whistler waves should propagate towards upstream and downstream regions from the rippled ramp. It is shown that the downstream pattern of the fields and ion distributions should follow the rippling pattern, while collisionless relaxation should be faster than in the stationary planar case.

Numerical Study on Infrared Radiation Characteristics of Dual-Engine Fighter Based on RMCM

The study of the infrared radiation characteristics of dual-engine fighters provides theoretical support for the stealth design of future fighters. This paper takes a dual-engine fighter as the research object and uses the Reverse Monte Carlo Method (RMCM) to carry out the numerical simulation of its infrared radiation characteristics. The results show that: in the 3-5μm band, with the Mach number increasing, the peak infrared radiation intensity increases, and the solid radiation accounts for about 90% of the total radiation at a detection angle of 180°; in the 8-14μm band, the infrared radiation intensity is distributed in the shape of "8". When the Mach number increases, the infrared radiation intensity of the abdomen of the fuselage is higher than that of the back.

An efficient multi-fidelity simulation method for adaptive cycle engine ejector nozzle performance evaluation

The three-bypass duct structure of Adaptive Cycle Engine (ACE) provides unique performance advantages while also challenging the design of its exhaust system. The separated exhaust nozzles widely adopted in current research may not meet the performance requirement of ACE. The feature of combining two exhaust streams to discharge makes the ejector nozzle suitable as the exhaust system of ACE. However, due to the connection between the third bypass duct and the ejector nozzle, the ejector nozzle and ACE form a highly coupled system, thus it is necessary to perform an integrated simulation of the ejector nozzle and ACE. The multi-fidelity simulation could offer accurate performance simulation as well as consider the complicated component matching working relations. In this paper, a multi-fidelity simulation method for ACE ejector nozzle simulation is developed for fast and accurate performance evaluation, and the design requirements of ACE ejector nozzle are initially explored. A multi-fidelity joint simulation method for the (0-D) ACE performance model and the (2-D) ejector nozzle Computational Fluid Dynamics (CFD) model is constructed through a dynamically optimized and updated surrogate model. The result shows, when ensuring the ejector nozzle does not affect the ACE working point, the peak gross thrust coefficient occurs in subsonic cruise condition as 0.9895. The gross thrust coefficients of acceleration conditions decrease with the increase of flight Mach number due to the increase of shroud divergent angle. Only 350 times CFD simulations are required to complete multi-fidelity simulations of 29 ACE operating points (As a comparison, the de-coupled method requires about 105 times CFD simulations for the whole characteristics, and the fully integrated method requires about 100 times CFD simulations for one ACE operating point). The diameter of the ejector shroud shoulder must be variable, otherwise it will cause a maximum thrust loss of 22.6% at dry acceleration condition and fan surge at cruise and reheat acceleration conditions. This method can further realize the evaluation of the installed performance and the control law optimal design of the ACE ejector nozzle.

Theoretical and numerical study of the binary scaling law for electron distribution in thermochemical non-equilibrium flows under extremely high Mach number

The binary scaling law is a classical similarity law used in analysing hypersonic flow fields. The objective of this study is to investigate the applicability of the binary scaling law in thermochemical non-equilibrium airflow. Dimensional analysis of vibrational and electron–electronic energy conservation equations was employed to explore the theoretical reasons for the failure of the binary scaling law. Numerical simulation based on a multi-temperature model (translational–rotational temperature T, electron–electronic excitation temperature ${T_e}$ and the vibrational temperatures of ${\textrm{O}_2}$ and ${\textrm{N}_2}$ , $\; {T_{{v_{{\textrm{O}_2}}}}}$ and ${T_{{v_{{\textrm{N}_2}}}}}$ ) with two chemical models (the Gupta model and the Park model) was adopted to study the accuracy of the binary scaling law for electron distribution at high altitude with extremely high Mach number. The results of theoretical analysis indicate that the three-body collision reactions and the translation–electron energy exchange from collisions between electrons and ions, ${Q_{t - e\_ions}}$ , can cause the failure of the binary scaling law. The results of numerical simulation show that the electron-impact ionization reactions are the main reasons for the invalidation of the binary scaling law for electron distribution at high altitude with high Mach number. With an increase of free-stream Mach number, the negative effect on the binary scaling law caused by ${Q_{t - e\_ions}}$ cannot be ignored.

Oscillation of the shock train under synchronous variation of incoming Mach number and backpressure

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.

An experimental and numerical investigation of the effects of the diaphragm pressure ratio and its position on a heated shock tube performance

AbstractThe main purpose of this work was to investigate the performance of heated shock tube (ST) with different pressure ratios and diaphragm positions numerically and experimentally. The numerical model was developed to simulate the fluid flow inside a shock tube test facility located at Qatar University. The shock tube was a cast‐iron hollow tube with 6 m length, 50 mm internal diameter and 10 mm thickness. ST driver and driven sections were filled with helium–argon mixture and air. The driven section was heated up to 150°C using coils. At the middle of the ST, the diaphragm was made of aluminium sheet (0.5 mm) layers. Five different pressure ratios were implemented during the experiment, and performance evaluation depended on the strength of the incident shock Mach number. The inviscid numerical model solver used transient two‐dimensional time‐accurate Navier–Stokes CFD. The model introduced a parametric study regarding three different diaphragm positions (1m, 2m and 3m) and five pressure ratios (6–10) for each position. In addition to yielding the incident and reflected wave Mach number, reflected wave temperature was also considered a shock tube performance indicator. The incident Mach numbers for the diaphragm middle position from the experiment were compared against those conducted from the model, and good matching was observed. The parametric study results showed that at high‐pressure ratios, diaphragm Positions 1 and 3 could generate a 7.4% increase in shock wave Mach number compared with the diaphragm position‐2 model. Moreover, the diaphragm position‐3 model tends to have a 2% increase in the temperature behind the reflected shock wave compared with the other two positions.

Determination of entropy-swallowing point of blunt hypersonic cone

Entropy swallowing is the phenomenon where the entropy layer is swallowed by the boundary layer developing downstream. In this work, the entropy-swallowing point over a 5° blunt hypersonic cone is investigated theoretically and numerically. An equation for obtaining the location of the entropy-swallowing point is derived herein that requires only the free-stream parameters. This equation also reveals the relationship between the entropy-swallowing point and the nose radius under the constant free-stream Mach number. The analytical results are entirely consistent with direct numerical simulation results. Furthermore, the influence of the Mach number and Reynolds number on the entropy-swallowing point is investigated. The location of the entropy-swallowing point depends more strongly on the Reynolds number than on the Mach number, with the entropy-swallowing point shifting further downstream as the Reynolds number increases. However, as the Mach number increases, its impact on the entropy-swallowing point decreases. Furthermore, the influence of the boundary-layer displacement also decreases with an increase of the Mach number.