There are great differences between ground test of a flight vehicle's aerodynamics and that of a dual-mode scramjet When ground test facilities are used to do experiments on a dual-mode scramjet engine, the components of incoming flow and its parameters, including total pressure, total temperature (or total enthalpy) and velocity must be simulated.

A wind tunnel is designed under asymmetric incoming flow and the compression fields in the isolator have been investigated using wall static and pitot pressure measurements, schlieren photography.

The performance of the compression ramp are studied mostly at the design point and off-design point in the condition of uniform supersonic incoming flow, and slightly at the design point in the condition of non-uniform incoming flow.

3. The effects of such elements as primary flow total pressure,flow rate,degree of fuel rich,nozzle structure,flying speed and secondary combustion on ejecting mode performance were researched.

Experimental and numerical methods were conducted in study the Dual-mode Scramjet(DMS) and the subsonic combustion was translated into supersonic combustion, at the simulated condition, which corresponds the flying state of Ma=6,H=25km.

The flow field and plasma sheath layer properties such as shock-wave shape, surface pressure, gas components, electron density and electronic collision frequency and so on are given in the range of free stream M∞ = 20-26, Re= (ρ U ,a)/μ= 1.5 ×103 -5.5×106 (where a is the radius of blunted nose).

The microorganisms in activated carbon effluent consisted mostly of heterotrophic bacillus and the total bacteria number was five times as high as that of the inflow, i.e.

Nitrogen removal of wetlands under 40 different inflow loadings were studied in the field during 15 months.

The outflow loading and total nitrogen (TN) removal rate of these beds under different inflow loadings and pollution loadings were investigated.

The inflow loadings of 4 subsurface flow systems (SFS) ranged from 400 to 8000 mg·(m2·d)-1, while outflow loadings were less than 7000 mg·(m2·d)-1.

The results showed that the inflow and outflow loading of TN removal rate in SFS presented an obvious linear relationship.

The nonsteady conditions in the incoming flow are characterized by the Strouhal number.

For a network with Poisson incoming flow of customers (particles) and unit time of the motion of servers (annihilators), we obtain the limit distribution of the number of customers at the node for a fixed general number of nodes.

For a network with Poisson incoming flow of customers (particles) and unit time of the motion of servers (annihilators), we obtain the limit distribution of the number of customers at the node for a fixed general number of nodes.

A numerical analysis reveals a nonattracting chaotic invariant set Λ that determines the scattering and trapping of particles from the incoming flow.

The hierarchy is found to have certain properties due to an infinite number of intersections of the stable manifold in Λ with a material line consisting of particles from the incoming flow.

A numerical calculative method of computing aerodynamic characteristics for sideslipping thin wings of arbitrary planform with dihedral in subsonic flow is presented in this paper. It extends the Vortex Lattics Method used for freestream flow that is paralled to the symmetric plane of the wing to sideslip with the aid of introducing a new coordinate system.Some concrete examples given in the paper were compared with other available methods and experimental values. The results of comparision are satisfied.

Experimental results of heat transfer to rough walls are given here for sphere cone models at Mach number 5. The nose radius of the models is 27.4mm and base diameter 60mm. Five models have been tested with different roughness in its bead diameter range from 0 to 0.9mm. The tests were conducted in a conventional hypersonic wind tunnel at total pressures from 10kg/cm2 and Reynolds numbers ReD from 0.8×106 to 3.6×106.The test results indicate that the smooth wall model heating is the laminar flow heating, its...

Experimental results of heat transfer to rough walls are given here for sphere cone models at Mach number 5. The nose radius of the models is 27.4mm and base diameter 60mm. Five models have been tested with different roughness in its bead diameter range from 0 to 0.9mm. The tests were conducted in a conventional hypersonic wind tunnel at total pressures from 10kg/cm2 and Reynolds numbers ReD from 0.8×106 to 3.6×106.The test results indicate that the smooth wall model heating is the laminar flow heating, its heat flux at the stagnation point is quite close to the theoretical data, and the influence of roughness at low total pressure (10kg/ cm2) occurs mainly to promote the transition and development of a boundary layer. With the increasing total pressure in the wind tunnel the static pressure on model and local Reynolds numbers increase correspondently. In this case the effect of roughness on heat transfer becomes remarkable, and the most remarkable region appears at sonic point region on the nose. At the highest total pressure (pt =45kg/cm2) and with the largest roughness (d = 0.9mm) the ratio of rough wall heat flux to laminar flow smooth one could be up to 4 except at stagnation point, where it could approach 6. Its raise seems to be related to the local shape change in the vicinity of the stagnation point.

An iterative method for solving two-dimensional compressible boundary layer equations and steady state equations of heat conduction simultaneously is presented, and also a FORTRAN program for calculation of the local convec-tive heat transfer coefficients over air-cooled vane surface by means of this method is provided. The approximate integral method is used for solving boun- dary layer equations end the finite element method is applied to calculating the steady temperature field of the blade.The input of the...

An iterative method for solving two-dimensional compressible boundary layer equations and steady state equations of heat conduction simultaneously is presented, and also a FORTRAN program for calculation of the local convec-tive heat transfer coefficients over air-cooled vane surface by means of this method is provided. The approximate integral method is used for solving boun- dary layer equations end the finite element method is applied to calculating the steady temperature field of the blade.The input of the program consists of geometry of the blade, pressure or velocity distribution of gas flow external to the boundary layer, entrance flow conditions, internal cooling conditions, nodal numbers and coordinates of the elements. The output includes all principal boundary layer parameters, such as heat transfer coefficients and temperature distribution on the surface, and temperature distribution inside the blade. A numerical example has been calculated and the results are favourably compared with the theoretical and experimental data given by other authors.