Effects of entry conditions on cracked kerosene-fueled supersonic combustor performance

Supersonic combustion of thermally cracked kerosene was experimentally investigated in two model supersonic combustors with different entry cross-section areas. Effects of entry static pressure, entry Mach number, combustor entry geometry, and injection scheme on combustor performance were systematically investigated and discussed based on the measured static pressure distribution and specific thrust increment due to combustion. In addition, the methodology for characterizing flow rate and composition of cracked kerosene was detailed. Using a pulsed Schlieren system, the interaction of supercritical and cracked kerosene jet plumes with a Mach 2.5 crossflow was also visualized at different injection temperatures. The present experimental results suggest that the use of a higher combustor entry Mach number as well as a larger combustor duct height would suppress the boundary layer separation near the combustor entrance and avoid the problem of inlet un- start.

A novel oblique detonation structure and its stability

Oblique detonation structures induced by the wedge in the supersonic combustible gas mixtures are simulated numerically. The results show that the stationary oblique detonation structures are influenced by the gas flow Mach number, and a novel critical oblique detonation structure, which is characterized by a more complicated wave system, appears in the low Mach number cases. By introducing the inflow disturbance, its nonstationary evolution process is illustrated and its stability is verified.

Numerical investigation of dynamic stall of an oscillating aerofoil

The flow structure around an NACA 0012 aerofoil oscillating in pitch around the quarter-chord is numerically investigated by solving the two-dimensional compressible N-S equations using a special matrix-splitting scheme. This scheme is of second-order accuracy in time and space and is computationally more efficient than the conventional flux-splitting scheme. A 'rigid' C-grid with 149 x 51 points is used for the computation of unsteady flow. The freestream Mach number varies from 0.2 to 0.6 and the Reynolds number from 5000 to 20,000. The reduced frequency equals 0.25-0.5. The basic flow structure of dynamic stall is described and the Reynolds number effect on dynamic stall is briefly discussed. The influence of the compressibility on dynamic stall is analysed in detail. Numerical results show that there is a significant influence of the compressibility on the formation and convection of the dynamic stall vortex. There is a certain influence of the Reynolds number on the flow structure. The average convection velocity of the dynamic stall vortex is approximately 0.348 times the freestream velocity.

Numerical-simulation of low supersonic-flow around a sphere

The controlled equations defined in a physical plane are changed into those in a computational plane with coordinate transformations suitable for different Mach number M(infinity). The computational area is limited in the body surface and in the vicinities of detached shock wave and sonic line. Thus the area can be greatly cut down when the shock wave moves away from the body surface as M(infinity) --> 1. Highly accurate, total variation diminishing (TVD) finite-difference schemes are used to calculate the low supersonic flowfield around a sphere. The stand-off distance, location of sonic line, etc. are well comparable with experimental data. The long pending problem concerning a flow passing a sphere at 1.3 greater-than-or-equal-to M(infinity) > 1 has been settled, and some new results on M(infinity) = 1.05 have been presented.

High-speed streams from coronal holes and the accelerating mechanism of the solar-wind

In this paper, the general Mach number equation is derived, and the influence of typical energy forms in the solar wind is analysed in detail. It shows that the accelerating process of the solar wind is influenced critically by the form of heating in the corona, and that the transonic mechanism is mainly the result of the adjustment of the variation of the crosssection of flowing tubes and the heat source term.The accelerating mechanism for both the high-speed stream from the coronal hole and the normal solar wind is similar. But, the temperature is low in the lower level of the coronal hole and more heat energy supply in the outside is required, hence the high speed of the solar wind; while the case with the ordinary coronal region is just the opposite, and the velocity of the solar wind is therefore lower. The accelerating process for various typical parameters is calculated, and it is found that the high-speed stream may reach 800 km/sec.

Starting Nozzle Flow Simulation Using K-G Two-Equation Turbulence Model

The starting process of two-dimensional nozzle flows has been simulated with Euler, laminar and k - g two-equation turbulence Navier-Stokes equations. The flow solver is based on a combination of LUSGS subiteration implicit method and five spatial discretized schemes, which are Roe, HLLE, MHLLE upwind schemes and AUSM+, AUSMPW schemes. In the paper, special attention is for the flow differences of the nozzle starting process obtained from different governing equations and different schemes. Two nozzle flows, previously investigated experimentally and numerically by other researchers, are chosen as our examples. The calculated results indicate the carbuncle phenomenon and unphysical oscillations appear more or less near a wall or behind strong shock wave except using HLLE scheme, and these unphysical phenomena become more seriously with the increase of Mach number. Comparing the turbulence calculation, inviscid solution cannot simulate the wall flow separation and the laminar solution shows some different flow characteristics in the regions of flow separation and near wall.

Numerical Simulation of R-M Instability

In order to capture shock waves and contact discontinuities in the field and easy to program with parallel computation a new algorithm is developed to solve the N-S equations for simulation of R-M instability problems. The method with group velocity control is used to suppress numerical oscillations, and an adaptive non-uniform mesh is used to get fine resolution. Numerical results for cylindrical shock-cylindrical interface interaction with a shock Mach number Ms=1.2 and Atwood number A=0.818, 0.961, 0.980 (the interior density of the interface/outer density p(1)/p(2) = 10, 50, 100, respectively), and for the planar shock-spherical interface interaction with Ms=1.2 and p(1)/p(2) = 14.28are presented. The effect of Atwood number and multi-mode initial perturbation on the R-M instability are studied. Multi-collisions of the reflected shock with the interface is a main reason of nonlinear development of the interface instability and formation of the spike-bubble structures In simulation with double mode perturbation vortex merging and second instability are found. After second instability the small vortex structures near the interface produced. It is important factor for turbulent mixing.

Experimental Studies on Scramjet Tested in a Freejet Facility

Two different type scramjet models with side-wall compression and top-wall compression inlets have been tested in HPTF (Hypersonic Propulsion Test Facility) under the experimental conditions of Mach number 5.8, total temperature 1700K, total pressure 4.5MPa and mass flow rate 3.5kg/s. The liquid kerosene was used as main fuel for the scramjets. In order to get fast ignition in the combustor, a small amount of hydrogen was used as a pilot. A strut with alternative tail was employed for increasing the compression ratio and for mixing enhancement in the side-wall compression case. Recessed cavities were used as a flameholder for combustion stability. The combustion efficiency was estimated by one dimensional theory. The uniformity of the facility nozzle flow was verified by a scanning pitot rake. The experimental results showed that the kerosene fuel was successfully ignited and stable combustion was achieved for both scramjet models. However the thrusts were still less than the model drags due to the low combustion efficiencies.

Numerical Investigation of Toroidal Shock Wave Focusing

In this paper,focusing of a toroidal shock wave propagating from a shock tube of an- nular cross-section into a cylindrical chamber was investigated numerically with the dispersion- controlled scheme. For CFD validation, the numerical code was rst applied to calculate both viscous and inviscid ows at a low Mach number of 1.5, which was compared with the experi- ment results and got better consistency. Then the validated code was used to calculate several cases for high Mach numbers. From the result, several major factors that in uent the ow, such as the Mach number and the viscosity, were analyzed detailedly and along with the high Mach number some unusual ow structure was observed and explained theoretically

Research progress on high-enthalpy and hypersonic flows

The research progress on high-enthalpy and hypersorlic flows having been achieved in the Institute of Mechanics, Chinese Academy of Sciences, is reported in this paper. The paper consists of three main parts: The first part is on the techniques to develop advanced hypersonic test facilities, in which the detonation-driven shock-reflected tunnel and the detonation-driven shock-expanded tube are introduced. The shock tunnel can be used for generating hypersonic flows of a Mach number ranging from 10 to 20, and the expansion tube is applicable to simulate the flows with a speed of 7 similar to 10km/s. The second part is dedicated to the shock tunnel nozzle flow diagnosis to examine properties of the hypersonic flows thus created. The third part is on experiments and numerical simulations. The experiments include measuring the aerodynamic pitching moment and heat transfer in hypersonic flows, and the numerical work reports nozzle flow simulations and flow non-equilibrium effects on the possible experiments that may be carried out on the above-mentioned hypersonic test facilities.

郭永怀文集

Experimental Demonstration Of A New Concept Of Drag Reduction And Thermal Protection For Hypersonic Vehicles

A new idea of drag reduction and thermal protection for hypersonic vehicles is proposed based on the combination of a physical spike and lateral jets for shock-reconstruction. The spike recasts the bow shock in front of a blunt body into a conical shock, and the lateral jets work to protect the spike tip from overheating and to push the conical shock away from the blunt body when a pitching angle exists during flight. Experiments are conducted in a hypersonic wind tunnel at a nominal Mach number of 6. It is demonstrated that the shock/shock interaction on the blunt body is avoided due to injection and the peak pressure at the reattachment point is reduced by 70% under a 4A degrees attack angle.

Direct numerical simulation of hypersonic boundary layer transition over a blunt cone with a small angle of attack

The direct numerical simulation of boundary layer transition over a 5° half-cone-angle blunt cone is performed. The free-stream Mach number is 6 and the angle of attack is 1°. Random wall blow-and-suction perturbations are used to trigger the transition. Different from the authors’ previous work [Li et al., AIAA J. 46, 2899(2008)], the whole boundary layer flow over the cone is simulated (while in the author’s previous work, only two 45° regions around the leeward and the windward sections are simulated). The transition location on the cone surface is determined through the rapid increase in skin fraction coefficient (Cf). The transition line on the cone surface shows a nonmonotonic curve and the transition is delayed in the range of 0° ≤ θ ≤ 30° (θ = 0° is the leeward section). The mechanism of the delayed transition is studied by using joint frequency spectrum analysis and linear stability theory (LST). It is shown that the growth rates of unstable waves of the second mode are suppressed in the range of 20° ≤ θ ≤ 30°, which leads to the delayed transition location. Very low frequency waves VLFWs� are found in the time series recorded just before the transition location, and the periodic times of VLFWs are about one order larger than those of ordinary Mack second mode waves. Band-pass filter is used to analyze the low frequency waves, and they are deemed as the effect of large scale nonlinear perturbations triggered by LST waves when they are strong enough.The direct numerical simulation of boundary layer transition over a 5° half-cone-angle blunt cone is performed. The free-stream Mach number is 6 and the angle of attack is 1°. Random wall blow-and-suction perturbations are used to trigger the transition. Different from the authors’ previous work [ Li et al., AIAA J. 46, 2899 (2008) ], the whole boundary layer flow over the cone is simulated (while in the author’s previous work, only two 45° regions around the leeward and the windward sections are simulated). The transition location on the cone surface is determined through the rapid increase in skin fraction coefficient (Cf). The transition line on the cone surface shows a nonmonotonic curve and the transition is delayed in the range of 20° ≤ θ ≤ 30° (θ = 0° is the leeward section). The mechanism of the delayed transition is studied by using joint frequency spectrum analysis and linear stability theory (LST). It is shown that the growth rates of unstable waves of the second mode are suppressed in the range of 20° ≤ θ ≤ 30°, which leads to the delayed transition location. Very low frequency waves (VLFWs) are found in the time series recorded just before the transition location, and the periodic times of VLFWs are about one order larger than those of ordinary Mack second mode waves. Band-pass filter is used to analyze the low frequency waves, and they are deemed as the effect of large scale nonlinear perturbations triggered by LST waves when they are strong enough.

Configuration optimization of hypersonic vehicles under transitional flow conditions

The rarefied gas effects on several configurations are investigated under hypersonic flow conditions using the direct simulation Mont Carlo method. It is found that the Knudsen number, the Mach number, and the angle of attack all play a mixed role in the aerodynamics of a flat plate. The ratio of lift to drag decreases as the Knudsen number increases. Studies on 3D delta wings show that the ratio of lift to drag could be increased by decreasing the wing thickness and/or by increasing the wing span. It is also found that the waveriders could produce larger ratio of lift to drag as compared with the delta wing having the same length, wing span, and cross section area.

Simulation of gas flow over a micro cylinder

Gas flow over a micro cylinder is simulated using both a compressible Navier-Stokes solver and a hybrid continuum /particle approach. The micro cylinder flow has low Reynolds number because of the small length scale and the low speed, which also indicates that the rarefied gas effect exists in the flow. A cylinder having a diameter of 20 microns is simulated under several flow conditions where the Reynolds number ranges from 2 to 50 and the Mach number varies from 0.1 to 0.8. It is found that the low Reynolds number flow can be compressible even when the Mach number is less than 0.3, and the drag coefficient of the cylinder increases when the Reynolds number decreases. The compressible effect will increase the pressure drag coefficient although the friction coefficient remains nearly unchanged. The rarefied gas effect will reduce both the friction and pressure drag coefficients, and the vortex in the flow may be shrunk or even disappear.