An experimental investigation on unsteady aerodynamic characteristics was carried out for two flat--plate delta wings of aspect ratio 1. 5 and 2. 0 undergoing large amplitude pitching motions inthe 3 × 2.5 low -- speed wind tunnel at NUAA.
A fully implicit Lower-Upper-Symmetric-Gauss-Seidel(LU-SGS) with pseudo time sub-iteration and the modified Jameson's central scheme are applied to solve thin layer Navier-Stokes(N-S) equations with laminar hypothesis and Baldwin-Lomax(B-L) model for the vortical flows around delta wings at high angles of attack(AOA).
The rolling-up of the leading-edge vortex sheet and interaction with the trailing-edge vortex are simulated numerically by using the two-dimensional unsteady analogue and vortex-in-cell method.
Effect of mach number on hypersonic flow past a delta wing with blunt edges
Supersonic three-dimensional flow around a delta wing with blunted leading edges
We consider the problem of steady flow of an inviscid, non-heat-conducting gas about a delta wing which is spherically blunted at the nose and cylindrlcally blunded on the leading edges, at an angle of attack.
Numerical solution of the problem of supersonic gas flow past an arbitrary delta wing surface in the compression region
We study the hypersonic flow of an inviscid ideal gas past a delta wing of small aspect ratio at a finite angle of attack.
Experimental study of flow about flat delta wings at m=5 and angles of attack from 0 to 70°
Results are presented of an experimental study of the flow about flat delta wings with sharp and blunt leading edges and sweep angles χ=60, 70, 80° for a freestream Mach number of 5 in the range of angles of attack from 0 to 70°.
Hypersonic flow past delta wings of a certain class at angle of attack with an attached shock
Supersonic aerodynamic characteristics of delta wings at high angles of attack
As examples we consider delta and double-delta wings.
Effect of mach number on hypersonic flow past a delta wing with blunt edges
Supersonic three-dimensional flow around a delta wing with blunted leading edges
We consider the problem of steady flow of an inviscid, non-heat-conducting gas about a delta wing which is spherically blunted at the nose and cylindrlcally blunded on the leading edges, at an angle of attack.
We study the hypersonic flow of an inviscid ideal gas past a delta wing of small aspect ratio at a finite angle of attack.
A delta wing with blunt edges at low angles of attack in hypersonic flow
This paper fully discusses the basic principle and experimental methods of determining aircraft longitudinal dynamic stability derivative during using half-model and various oscillations-free oscillation and fored oscillation. It briefly describes the self-designed and self-menufactured dynamic balance. According to the oomperison of the preliminary experimental results of delta wing model in our 250×200mm2 high speed wind tunnel with FFA62 experimental results, it is proved that this balance is useful. But...
This paper fully discusses the basic principle and experimental methods of determining aircraft longitudinal dynamic stability derivative during using half-model and various oscillations-free oscillation and fored oscillation. It briefly describes the self-designed and self-menufactured dynamic balance. According to the oomperison of the preliminary experimental results of delta wing model in our 250×200mm2 high speed wind tunnel with FFA62 experimental results, it is proved that this balance is useful. But in the course of our experiments, many problems occurred, hence the last part of this paper offers some suggestions as to the solution of them and states how we should endeavour later on.
The three-dimensional unsteady second-order non-homogeneous differential equation has been derived by superposition of a small disturbance on a given steady three-dimensional flow. Based on the assumption of high Mach numbers this second-order equation for unsteady flow reduces to a form analogous to that for steady flow. This makes it possible to solve the equation by methods used in the second-order theory for steady flows. In the course of solution the flows are constrained and corrected according to the...
The three-dimensional unsteady second-order non-homogeneous differential equation has been derived by superposition of a small disturbance on a given steady three-dimensional flow. Based on the assumption of high Mach numbers this second-order equation for unsteady flow reduces to a form analogous to that for steady flow. This makes it possible to solve the equation by methods used in the second-order theory for steady flows. In the course of solution the flows are constrained and corrected according to the PLK method, and singularities are thus eliminated. The crucial point in this procedure is to find the correct particular solutions. Two particular solutions are used. One is the approximate three-dimensional particular solution. The other is obtained under the assumption of local two-dimensionality. In addition, the uniform particular solution is given, from which the uniform second-order solutions may be obtained. Then, we have treated the unsteady problem for delta wings with low aspect ratio and supersonic leading edges. The Mach number range for application of the present theory is from supersonic to low hypersonic values with reduced frequencies up to near unity. The theoretical results derived in this work can be used to calculate the unsteady aerodynamic characteristics of wings having arbitrary airfoil sections.As experimental information for similar conditions is not yet available, we can only compare our results with those of Liu D. D. . For this reason, only the derivation for a flat delta wing oscillating at low frequencies has been carried out and an analytical expression is obtained for the first order expansion of the unsteady velocity potential. In the range of Mach numbers 4 to 8, our results are in agreement with those of Lui D. D. .It is also shown that under conditions of three-dimensional thin wings the second-order theory is valid up to Mδ=1.0, while according to application of the second-order theory to bodies of revolution by Van Dyke, the useful upper limit of M5 is only 0.7. Hence, with Mδ=0.7-1 .0, the principal non-linear effects can be calculated by our second-order theory, while for thin wings the third-order terms connected with heat transfer and entropy change can be ignored.
Velocity gradient,pressure gradient and circulation gradient are used to model the separated vortex flow field over a leading edge delta wing.The deceleration or adversed pressure of out-flow in the direction of x-axis promotes the breakdown of the vortex core,so does the larger circulation gradient in the direction of the axis.The induced larger adversed pressure gradient of the vortex core in the direction of axis or larger radial pressure difference results in the vortex breakdown,so the idea is more completely...
Velocity gradient,pressure gradient and circulation gradient are used to model the separated vortex flow field over a leading edge delta wing.The deceleration or adversed pressure of out-flow in the direction of x-axis promotes the breakdown of the vortex core,so does the larger circulation gradient in the direction of the axis.The induced larger adversed pressure gradient of the vortex core in the direction of axis or larger radial pressure difference results in the vortex breakdown,so the idea is more completely obtained about the vortex breakdown mechanism.The axisymmetric vortex breakdown can not take place until the Reynolds number becomes larger.
三角翼
delta wing
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