Unsteady Aerodynamic Analysis of an Oscillating Cascade at Large Incidence

1999 ◽  
Author(s):  
Jong-Shang Liu ◽  
Durbha V. Murthy
Keyword(s):  
Mach Number ◽  
Fluid Dynamic ◽  
Leading Edge ◽  
Two Dimensional ◽  
Edge Region ◽  
Unsteady Aerodynamic ◽  
High Incidence ◽  
Flutter Analysis ◽  
High Mach ◽  
Good Agreement

The flutter analysis capability of the quasi-3D aeroelastic computational fluid dynamic (CFD) code UNSFLO is evaluated by comparing to unsteady pressure results in an oscillating cascade. The geometry is two-dimensional and the oscillation is well-controlled. Time unsteady UNSFLO results are compared with data at inlet Mach number from 0.2 to 0.8 with incidence 0° and 10°. Three reduced frequencies, 0.4, 0.8, and 1.2 with inter-blade phase angle of 180 degrees were tested. The calculated steady state loadings show good agreement with data at zero incidence. The correlations become worse for high incidence angles because of the separation. The calculated aerodynamic work capture the chordwise distribution except in the near leading edge region. The correlations become also worse for high incidence from the leading edge to midchord especially at high Mach number.

An experimental study of turbulent separating and reattaching flows at a high Mach number

1972 ◽  
Vol 52 (3) ◽  
pp. 425-435 ◽  
Author(s):  
J. P. Batham
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Leading Edge ◽  
Reynolds Numbers ◽  
Shear Layers ◽  
Two Dimensional ◽  
Pressure Coefficients ◽  
High Mach ◽  
Incipient Separation ◽  
Good Agreement

Separating and reattaching flows in a two-dimensional compression corner were investigated experimentally at a Mach number of 7·0 and Reynolds numbers (based on the distance from the leading edge to the corner) of 4·75 × 106, 9·51 × 106 and 1·55 × 107. Heat-transfer measurements and Pitot traverses in the upstream boundary layer showed that the boundary layer had become fully turbulent at the start of the interactions. Increases in the Reynolds number gave increases in the length of separated shear layers and decreases in the corner angle required for incipient, separation. The reattachment pressure coefficients gave good agreement with the criterion of Batham (1969).

A Navier-Stokes Analysis of the Stall Flutter Characteristics of the Buffum Cascade

2000 ◽  
Author(s):  
Stefan Weber ◽  
Max F. Platzer
Keyword(s):  
Mach Number ◽  
Turbulence Modeling ◽  
Separation Bubble ◽  
Leading Edge ◽  
Navier Stokes ◽  
Implicit Time ◽  
Fully Implicit ◽  
High Incidence ◽  
Stall Flutter ◽  
Good Agreement

Numerical stall flutter prediction methods are highly needed as modern jet engines require blade designs close to the stability boundaries of the performance map. A Quasi-3D Navier-Stokes code is used to analyze the flow over the oscillating cascade designed and manufactured by Pratt & Whitney, and studied at the NASA Glenn Research Center by Buffum et al. The numerical method solves for the governing equations with a fully implicit time-marching technique in a single passage by making use of a direct-store, periodic boundary condition. For turbulence modeling the Baldwin-Lomax model is used. To account for transition, the criterion to predict the onset location suggested by Baldwin and Lomax is incorporated. Buffum et al. investigated two incidence cases for three different Mach numbers. The low-incidence case at a Mach number of 0.5 exhibited the formation of small separation bubbles at reduced oscillation frequencies of 0.8 and 1.2. For this case the present approach yielded good agreement with the steady and oscillatory measurements. At high-incidence at the same Mach number of 0.5 the measured steady-state pressure distribution and the separation bubble on the upper surface was also found in good agreement with the experiment. But computations for oscillations at high-incidence failed to predict the negative damping contribution caused by the leading edge separation.

A Navier–Stokes Analysis of the Stall Flutter Characteristics of the Buffum Cascade

2000 ◽  
Vol 122 (4) ◽  
pp. 769-776 ◽  
Author(s):  
Stefan Weber ◽  
Max F. Platzer
Keyword(s):  
Mach Number ◽  
Turbulence Modeling ◽  
Separation Bubble ◽  
Leading Edge ◽  
Navier Stokes ◽  
Implicit Time ◽  
Fully Implicit ◽  
High Incidence ◽  
Stall Flutter ◽  
Good Agreement

Numerical stall flutter prediction methods are much needed, as modern jet engines require blade designs close to the stability boundaries of the performance map. A Quasi-3D Navier–Stokes code is used to analyze the flow over the oscillating cascade designed and manufactured by Pratt & Whitney, and studied at the NASA Glenn Research Center by Buffum et al. The numerical method solves for the governing equations with a fully implicit time-marching technique in a single passage by making use of a direct-store, periodic boundary condition. For turbulence modeling, the Baldwin–Lomax model is used. To account for transition, the criterion to predict the onset location suggested by Baldwin and Lomax is incorporated. Buffum et al. investigated two incidence cases for three different Mach numbers. The low-incidence case at a Mach number of 0.5 exhibited the formation of small separation bubbles at reduced oscillation frequencies of 0.8 and 1.2. For this case the present approach yielded good agreement with the steady and oscillatory measurements. At high incidence at the same Mach number of 0.5 the measured steady-state pressure distribution and the separation bubble on the upper surface was also found in good agreement with the experiment. But computations for oscillations at high incidence failed to predict the negative damping contribution caused by the leading edge separation. [S0889-504X(00)01304-0]

Modeling of External Pressure Gradient Influence on the Stability of Disturbances in the Boundary Layers of Compressed Gas

2013 ◽  
Vol 8 (4) ◽  
pp. 64-75
Author(s):  
Sergey Gaponov ◽  
Natalya Terekhova
Keyword(s):  
Mach Number ◽  
Pressure Gradient ◽  
Boundary Layers ◽  
Leading Edge ◽  
External Pressure ◽  
Transition To Turbulence ◽  
Compressed Gas ◽  
High Mach ◽  
The Stability ◽  
Very High

This work continues the research on modeling of passive methods of management of flow regimes in the boundary layers of compressed gas. Authors consider the influence of pressure gradient on the evolution of perturbations of different nature. For low Mach number M = 2 increase in pressure contributes to an earlier transition of laminar to turbulent flow, and, on the contrary, drop in the pressure leads to a prolongation of the transition to turbulence. For high Mach number M = 5.35 found that the acoustic disturbances exhibit a very high dependence on the sign and magnitude of the external gradient, with a favorable gradient of the critical Reynolds number becomes smaller than the vortex disturbances, and at worst – boundary layer is destabilized directly on the leading edge

scholarly journals Unsteady Pressure Measurements on a Biconvex Airfoil in a Transonic Oscillating Cascade

1986 ◽  
Vol 108 (1) ◽  
pp. 53-59 ◽  
Author(s):  
L. M. Shaw ◽  
D. R. Boldman ◽  
A. E. Buggele ◽  
D. H. Buffum
Keyword(s):  
Mach Number ◽  
Incidence Angle ◽  
Dynamic Pressure ◽  
Leading Edge ◽  
Pressure Transducers ◽  
High Incidence ◽  
Shock Motion ◽  
Plate Analysis ◽  
Phase Angles ◽  
Unsteady Pressure Measurements

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfoils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence angle of 0.0 deg and a Mach number of 0.65 in order to obtain results in a shock-free compressible flow field. Subsequent tests were performed at an angle of attack of 7.0 deg and a Mach number of 0.80 in order to observe the surface pressure response with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and −90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semichord reduced frequencies up to about 0.5 at a Mach number of 0.80). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Application of a Two-Dimensional Supersonic Passage Analysis to the Design of Compressor Rotors

1974 ◽  
Author(s):  
R. G. Hantman ◽  
A. A. Mikolajczak ◽  
F. J. Camarata
Keyword(s):  
Leading Edge ◽  
Inviscid Flow ◽  
Two Dimensional ◽  
Compressor Rotor ◽  
Pressure Distributions ◽  
Displacement Thickness ◽  
Blade Section ◽  
Comparison Of The Results ◽  
Rotational Method ◽  
Good Agreement

A description of a two-dimensional supersonic cascade passage analysis and its application to the design of a high hub-to-tip ratio supersonic compressor rotor is presented. The analysis, applicable to the case in which the inviscid flow is everywhere supersonic, includes an entrance region calculation which accounts for blade leading edge bluntness effects, and a passage and wake region calculation. The inviscid part of the analysis is solved using a rotational method of characteristics. The effect of the blade boundary layer displacement thickness is taken into consideration. Comparison of the results of the analysis with supersonic cascade data is made, showing good agreement in overall performance prediction, in blade surface static pressure distributions, and in achievement of the desired shock wave patterns. A comparison of the results of the analysis is made also with the performance of a blade section of a high hub-to-tip ratio supersonic compressor and acceptable agreement obtained.

Acoustic destablilization of vortices

1979 ◽  
Vol 290 (1372) ◽  
pp. 353-371 ◽  
Keyword(s):  
Boundary Condition ◽  
Growth Rate ◽  
Mach Number ◽  
Acoustic Waves ◽  
Polar Angle ◽  
Inviscid Flow ◽  
Two Dimensional ◽  
Low Mach Number ◽  
The Stability ◽  
Good Agreement

A line vortex which has uniform vorticity 2Ω 0 in its core is subjected to a small two-dimensional disturbance whose dependence on polar angle is e imθ . The stability is examined according to the equations of compressible, inviscid flow in a homentropic medium. The boundary condition at infinity is that of outgoing acoustic waves, and it is found that this capacity to radiate leads to a slow instability by comparison with the corresponding incompressible vortex which is stable. Numerical eigenvalues are computed as functions of the mode number m and the Mach number M based on the circumferential speed of the vortex. These are compared with an asymptotic analysis for the m = 2 mode at low Mach number in which it is found that the growth rate is (π/ 32) M 4 Ω 0 in good agreement with the numerical results.

A Simulation of a Supersonic Intake on a Hybrid Computer

1980 ◽  
Vol 31 (3) ◽  
pp. 197-220
Author(s):  
D.E. Colbourne
Keyword(s):  
Steady State ◽  
Mach Number ◽  
Dynamic Performance ◽  
Equations Of Motion ◽  
Two Dimensional ◽  
Frequency Responses ◽  
Lumped Parameter ◽  
Time Simulation ◽  
Hybrid Computer ◽  
High Mach

SummaryA simulation of the steady-state and dynamic performance of a supersonic powerplant intake is described. A lumped-parameter form of the equations of motion for a compressible fluid is developed. The resulting equations, together with steady-state characteristics, are used to model the performance of a two-dimensional supersonic intake. The model is implemented on the NGTE hybrid computer. The performance, both steady-state and transient, of the resulting real-time simulation has been investigated. Step and frequency responses at two high Mach number flight conditions have been examined and compared with test data. The favourable comparisons obtained demonstrate the validity of the chosen technique.

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