Separated flow

A brief overview of flow separation phenomena is provided. Langley has many active research programs in flow separation related areas. Three cases are presented which describe specific examples of flow separation research. In each example, a description of the fundamental fluid physics and the complexity of the flow field is presented along with a method of either reducing or controlling the extent of separation. The following examples are discussed: flow over a smooth surface with an adverse pressure gradient; flow over a surface with a geometric discontinuity; and flow with shock-boundary layer interactions. These results will show that improvements are being made in the understanding of flow separation and its control

Separated flow over bodies of revolution using an unsteady discrete-vorticity cross wake. Part 2: Computer program description

A method is developed to determine the flow field of a body of revolution in separated flow. The computer was used to integrate various solutions and solution properties of the sub-flow fields which made up the entire flow field without resorting to a finite difference solution to the complete Navier-Stokes equations. The technique entails the use of the unsteady cross flow analogy and a new solution to the two-dimensional unsteady separated flow problem based upon an unsteady, discrete-vorticity wake. Data for the forces and moments on aerodynamic bodies at low speeds and high angle of attack (outside the range of linear inviscid theories) such that the flow is substantially separated are produced which compare well with experimental data. In addition, three dimensional steady separated regions and wake vortex patterns are determined. The computer program developed to perform the numerical calculations is described

Structure of Evaporating and Combusting Sprays: Measurements and Predictions

Complete measurements of the structure of nonevaporating, evaporating and combusting sprays for sufficiently well defined boundary conditions to allow evaluation of models of these processes were obtained. The development of rational design methods for aircraft combustion chambers and other devices involving spray combustion were investigated. Three methods for treating the discrete phase are being considered: a locally homogeneous flow (LHF) model, a deterministic separated flow (DSF) model, and a stochastic separated flow (SSF) model. The main properties of these models are summarized

Flow separation in shock wave boundary layer interactions at hypersonic speeds

An assessment is presented for the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consists mainly of two dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the subsequent changes in the flow empirical correlations for incipient separation are reviewed. The effects of the Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed. The pertinent experimental data for the separated flow characteristics in separated turbulent boundary layer shock interaction are also presented and discussed

A steady separated viscous corner flow

An example is presented of a separated flow in an unbounded domain in which, as the Reynolds number becomes large, the separated region remains of size 0(1) and tends to a non-trivial Prandtl-Batchelor flow. The multigrid method is used to obtain rapid convergence to the solution of the discretized Navier-Stokes equations at Reynolds numbers of up to 5000. Extremely fine grids and tests of an integral property of the flow ensure accuracy. The flow exhibits the separation of a boundary layer with ensuing formation of a downstream eddy and reattachment of a free shear layer. The asymptotic (’triple deck’) theory of laminar separation from a leading edge, due to Sychev (1979), is clarified and compared to the numerical solutions. Much better qualitative agreement is obtained than has been reported previously. Together with a plausible choice of two free parameters, the data can be extrapolated to infinite Reynolds number, giving quantitative agreement with triple-deck theory with errors of 20% or less. The development of a region of constant vorticity is observed in the downstream eddy, and the global infinite-Reynolds-number limit is a Prandtl-Batchelor flow; however, when the plate is stationary, the occurrence of secondary separation suggests that the limiting flow contains an infinite sequence of eddies behind the separation point. Secondary separation can be averted by driving the plate, and in this case the limit is a single-vortex Prandtl-Batchelor flow of the type found by Moore, Saffman & Tanveer (1988); detailed, encouraging comparisons are made to the vortex-sheet strength and position. Altering the boundary condition on the plate gives viscous eddies that approximate different members of the family of inviscid solutions

The calculation of separated flow at helicopter bodies

For abstract see A81-47555

An improved analytical model of the separation region on boattail nozzles at subsonic speeds

A practical engineering calculation was developed to model the viscous effects of a separated, reverse flow region on afterbody pressures and drag. This viscous calculation was iteratively coupled with an inviscid flow calculation by means of an aerodynamic interface. A standard boundary layer displacement thickness was used to modify the afterbody shape where the flow was attached. A discriminating streamline calculation was developed to account for displacement effects of the reverse flow in separated regions with and without a flowing jet. The viscous flow calculation was coupled with a potential flow calculation. The analysis accurately predicted afterbody pressures and drag with variations in Reynolds number, Mach number, and afterbody shape

A numerical study on aerodynamic resonance in transonic separated flow

An ongoing numerical investigation of unsteady shock/boundary layer interaction on a 2-d supercritical airfoil in transonic flow is presented. Initially, the finitevolume URANS solver DLR-TAU is used to simulate self-sustained periodic shock oscillations well known as shock buffet. Next, emphasis is put on the fixed-point stability of the steady flow field below the shock buffet onset. Therefore the flow is perturbed in time with small sinusoidal deflections of the airfoil geometry and random impulses. With increasing angle of attack the mean flow is shown to develop a damped aerodynamic resonance, that degenerates finally towards self-amplification. The occurrence of the aerodynamic resonance is closely related to the development of shock-induced separation, accompanied by quasi-steady inverse shock motion

On the flow processes in sharply inclined and stalled airfoils in parallel movement and rotation

The purpose of this study is to obtain a deeper insight into the complicated flow processes on airfoils in the region of the buoyancy maxima. To this end calculated and experimental investigations are carried out on a straight stationary, a twisted stationary and a straight rotating rectangular wing. According to the available results the method gives results which can be applied sufficiently for flow applied firmly on all sides for all rotation values. The reliability of the method may be questioned for a flow undergoing transition from the attached to the separated state or for totally separated flow and higher rotation values

Aerodynamic noise research support Quarterly progress report, Jan. - Mar. 1966

Aerodynamic noise research - acoustic environment due to separated flows and oscillating shocks and flow visualization experiments with supersonic separated turbulent flo