Design Method for Subsonic and Transonic Cascade With Prescribed Mach Number Distribution

1992 ◽  
Vol 114 (3) ◽  
pp. 553-560 ◽  
Author(s):  
O. Le´onard ◽  
R. A. Van den Braembussche
Keyword(s):  
Mach Number ◽  
Flow Field ◽  
Pressure Distribution ◽  
Velocity Component ◽  
Design Method ◽  
Normal Velocity ◽  
Flow Solver ◽  
Number Distribution ◽  
Mach Number Distribution ◽  
Transonic Cascade

A iterative procedure for blade design, using a time marching procedure to solve the unsteady Euler equations in the blade-to-blade plane, is presented. A flow solver, which performs the analysis of the flow field for a given geometry, is transformed into a design method. This is done by replacing the classical slip condition (no normal velocity component) by other boundary conditions, in such a way that the required pressure or Mach number distribution may be imposed directly on the blade. The unknowns are calculated on the blade wall using the so-called compatibility relations. Since the blade shape is not compatible with the required pressure distribution, a nonzero velocity component normal to the blade wall evolves from the new flow calculation. The blade geometry is then modified by resetting the wall parallel to the new flow field, using a transpiration technique, and the procedure is repeated until the calculated pressure distribution has converged to the required one. Examples for both subsonic and transonic flows are presented and show a rapid convergence to the geometry required for the desired Mach number distribution. An important advantage of the present method is the possibility to use the same code for the design and the analysis of a blade.

Design Method for Subsonic and Transonic Cascade With Prescribed Mach Number Distribution

1991 ◽  
Author(s):  
O. Léonard ◽  
R. A. Van Den Braembussche
Keyword(s):  
Mach Number ◽  
Flow Field ◽  
Pressure Distribution ◽  
Velocity Component ◽  
Design Method ◽  
Normal Velocity ◽  
Flow Solver ◽  
Number Distribution ◽  
Mach Number Distribution ◽  
Transonic Cascade

An iterative procedure for blade design, using a Time Marching procedure to solve the unsteady Euler equations in the blade-to-blade plane is presented. A flow solver, which performs the analysis of the flow field for a given geometry, is transformed into a design method. This is done by replacing the classical slip condition (no normal velocity component) by other boundary conditions, in such a way that the required pressure or Mach number distribution may be imposed directly on the blade. The unknowns are calculated on the blade wall using the so-called compatibility relations. Since the blade shape is not compatible with the required pressure distribution, a non-zero velocity component normal to the blade wall evolves from the new flow calculation. The blade geometry is then modified by resetting the wall parallel to the new flow field, using a transpiration technique, and the procedure is repeated until the calculated pressure distribution has converged to the required one. Examples for both subsonic and transonic flows are presented and show a rapid convergence to the geometry required for the desired Mach number distribution. An important advantage of the present method is the possibility to use the same code for the design and the analysis of a blade.

On the Use of Direct and Inverse Numerical Flow Calculations for Supersonic Turbine Design

1994 ◽  
Author(s):  
F. Pommel
Keyword(s):  
Flow Field ◽  
Euler Equations ◽  
Velocity Component ◽  
Normal Velocity ◽  
Blade Design ◽  
Blade Profile ◽  
Flow Solver ◽  
Time Marching ◽  
Turbine Design ◽  
Marching Method

A procedure for blade design, using a time marching method to solve the Euler equations in the blade-to-blade plane is presented. This procedure uses an Office Nationale d’Etude et de Recherches Aeronautique flow solver. The classical slip conditions (no normal velocity component along the blade profile) has been replaced by another boundary conditions in such a way that the required pressure may be imposed directly. The orignal direct code was therefore transformed into an inverse solver. The unknows are calculated on the blade wall using the so-called compatibility relations. The blade geometry is then modified by resetting the wall parallel to the new flow field. The results obtained with this design process for a supersonic turbine blade of a space turbopump is presented.

Investigation of Wall Effect on Flow Characteristics in Supersonic Rarefied Gas Flow Past a Confined Square Cylinder Using the DSMC Method

2011 ◽  
Author(s):  
Mohamad M. Joneidipour ◽  
Reza Kamali
Keyword(s):  
Mach Number ◽  
Flow Field ◽  
Gas Flow ◽  
Rarefied Gas ◽  
Blockage Ratio ◽  
Parallel Plates ◽  
Square Cylinder ◽  
Rarefied Gas Flow ◽  
Number Distribution ◽  
Mach Number Distribution

In the present study, the effect of wall in supersonic rarefied gas flow past a square cylinder is numerically studied. Therefore, a supersonic rarefied gas flow over a square cylinder is simulated first. Then, the simulations are repeated for a square cylinder confined between two parallel plates. In both cases, the Mach number distribution in the flow field of supersonic rarefied gas over the square cylinder is obtained using the direct simulation Monte Carlo method. Close inspection of contour lines of Mach number over the square cylinder shows that a discontinuity in the flow field occurs across the shock wave at the slip regime while there is no discontinuity at the transition flow regime. In the present paper, the effect of blockage ratio (defined as the distance between two parallel plates divided by the cylinder length) on the Mach number distribution in the flow field of supersonic rarefied gas over the square cylinder is also studied. Meanwhile, the obtained results from both mentioned cases are compared to each other. It is found that the deviation of two sets of data diminishes gradually as the blockage ratio increases.

scholarly journals A preliminary study on static pressure probe mobile measurement method for flow field in transonic wind tunnel

2021 ◽  
Vol 39 (4) ◽  
pp. 786-793
Author(s):  
Haijun Deng ◽  
Bo Xiong ◽  
Xinfu Luo ◽  
Shaozun Hong ◽  
Qi Liu ◽  
...  
Keyword(s):  
Mach Number ◽  
Wind Tunnel ◽  
Flow Field ◽  
Static Pressure ◽  
Wind Tunnel Test ◽  
Flow Characteristics ◽  
Pressure Probe ◽  
Number Distribution ◽  
Mach Number Distribution ◽  
Transonic Wind Tunnel

The axial Mach number distribution of the core flow for model in a transonic wind tunnel is an important index to evaluate the performance of the flow field, which is usually measured by the centerline probe. In order to simulate the incoming flow characteristics without interference, the probe will extend from the support section to the shrinkage section, so the probe usually must has longer inches, more static pressure measuring points and smaller blockage requirements. In order to study the influence of the points of the centerline probe on the uniformity distribution of flow field, a new static pressure probe is designed, which is smaller and shorter than the centerline probe. On the basis of the stability of the flow field, the Mach number distribution of the flow field measured by the static pressure probe which is driven by the moving measuring mechanism. The characteristics of the measured values are studied by wind tunnel test. The results show that: when Ma ≤ 0.95, the overall distribution and value of Mach number obtained by the static pressure probe is basically the same as those obtained by the centerline probe, but some flow field details, which mainly shows that Mach number of the static pressure probe has smaller fluctuation, higher accuracy and better uniformity index.

Mach number distribution over the axis of supersonic underexpanded jets

1968 ◽  
Vol 15 (6) ◽  
pp. 1153-1157 ◽  
Author(s):  
Yu. P. Finat'ev ◽  
L. A. Shcherbakov ◽  
N. M. Gorskaya
Keyword(s):  
Mach Number ◽  
Underexpanded Jets ◽  
Number Distribution ◽  
Mach Number Distribution

Investigation of the Influence of Trailing Edge Shock Waves on Film Cooling Performance of Gas Turbine Airfoils

2007 ◽  
Author(s):  
M. Ochs ◽  
A. Schulz ◽  
H.-J. Bauer
Keyword(s):  
Mach Number ◽  
Shock Waves ◽  
Film Cooling ◽  
Suction Side ◽  
Test Rig ◽  
Local Cooling ◽  
Trailing Edge ◽  
Number Distribution ◽  
Mach Number Distribution ◽  
Turbine Airfoils

Transonic turbine stage flows are strongly influenced by shock waves. The oblique trailing edge shock generated at the pressure side impinges on the suction side of the neighboring airfoil leading to a significant alteration of the Mach number distribution. On film cooled turbine airfoils this shock interacts with the local cooling film. The present study deals with the investigation of this kind of shock wave – film cooling interaction. Experiments are conducted in a high pressure high temperature transonic test rig which allows setting engine realistic Reynolds numbers and Mach numbers, as well as temperature and density ratios. The generic test rig simulates a transonic region of an airfoil passage with the advantage of accessibility for optical measurement techniques. Coolant is ejected from a row of 5 cylindrical and 5 fanshaped holes at different locations relative to the position of shock impingement. Blowing ratios are varied within a range of 0.25<M<1.5. A simulated suction side Mach number distribution is generated with a Mach number Mam = 1.45 upstream and Mam = 1.14 downstream of the shock. Experimental data presented comprise spatially resolved and laterally averaged film cooling effectiveness and heat transfer coefficients within the vicinity of the interaction zone.

Calculation of Cascade Profiles From the Velocity Distribution

1974 ◽  
Vol 96 (4) ◽  
pp. 407-412 ◽  
Author(s):  
C. Lecomte
Keyword(s):  
Mach Number ◽  
Velocity Distribution ◽  
Subsonic Flow ◽  
A Priori ◽  
External Flow ◽  
Analytical Function ◽  
Graphic Display ◽  
Analytical Functions ◽  
Number Distribution ◽  
Mach Number Distribution

The method using the properties of analytical functions is applied to a plane, steady, inviscid, everywhere subsonic flow. From data fixed a priori concerning the external flow and some details of the profile, the hodograph is obtained as an analytical function whose real part is known on a contour. The set of imposed conditions being in general superabundant, the proposed Mach number distribution is corrected by means of a function whose form is fixed a priori, or rejected altogether. The problem is treated on a graphic display console connected with a computer, which provides also the profile corresponding to the calculated hodograph.

On the Design Criteria for Suppression of Secondary Flows in Centrifugal and Mixed Flow Impellers

1998 ◽  
Vol 120 (4) ◽  
pp. 723-735 ◽  
Author(s):  
M. Zangeneh ◽  
A. Goto ◽  
H. Harada
Keyword(s):  
Flow Field ◽  
Pressure Distribution ◽  
Centrifugal Compressor ◽  
Design Method ◽  
Three Dimensional ◽  
Design Guidelines ◽  
Secondary Flows ◽  
Mixed Flow ◽  
Optimum Pressure ◽  
Specific Speed

In this paper, for the first time, a set of guidelines is presented for the systematic design of mixed flow and centrifugal compressors and pumps with suppressed secondary flows and a uniform exit flow field. The paper describes the shape of the optimum pressure distribution for the suppression of secondary flows in the impeller with reference to classical secondary flow theory. The feasibility of achieving this pressure distribution is then demonstrated by deriving guidelines for the design specifications of a three-dimensional inverse design method, in which the blades are designed subject to a specified circulation distribution or 2πrVθ. The guidelines will define the optimum choice of the blade loading or ∂rVθ/∂m and the stacking condition for the blades. These guidelines are then used in the design of three different low specific speed centrifugal pump impellers and a high specific speed industrial centrifugal compressor impellers. The flows through all the designed impellers are computed numerically by a three-dimensional viscous code and the resulting flow field is compared to that obtained in the corresponding conventional impeller. The results show consistent suppression of secondary flows in all cases. The design guidelines are validated experimentally by comparing the performance of the inverse designed centrifugal compressor impeller with the corresponding conventional impeller. The overall performance of the stage with the inverse designed impeller with suppressed secondary flows was found to be 5 percent higher than the conventional impeller at the peak efficiency point. Exit flow traverse results at the impeller exit indicate a more uniform exit flow than that measured at the exit from the conventional impeller.

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