Aerodynamic and Heat Transfer Studies on HUB sections of a high pressure turbine blade: summary report

The stator and rotor blade hub sections designed for a high pressure turbine stage were studied in detail for their aerodynamic and heat transfer characteristics . The profile sections were tested in the National Aeronautical Laboratory Cascade Tunnels over a range of exit flow Mach numbers . The flow field and heat transfer characteristics of the cascades were also code based on Denton's method and the boundary layer code incorporating K- E turbulence model. The results indicated that there was a scope for improving the blade profile sections for high Mach number applications

ESTIMATION OF TURBOMACHINERY FLOW LOSSES THROUGH CASCADE TESTING

An invited talk titled "ESTIMATION OF TURBOMACHINERY FLOW LOSSES THROUGH CASCADE TESTING" was delivered as a part of a two day workshop on "LOSS MECHANISMS IN STEAM AND GAS TURBINES" held at M.S.Ramiah School of Advanced Studies, Bangalore, India. The presentation material used for the lecture is provided here

CASCADE TESTS ON CE 20 LOX TURBINE STATOR BLADE PROFILE

LPSC, ISRO as part of their CE 20 cryogenic engine development program, desired to get the liquid oxygen [LOX] turbo pump turbine profiles tested in the NAL Transonic Cascade Tunnel to obtain basic aerodynamic performance data. The LOX turbine consists of a nozzle, first rotor, stator and second rotor. The cascade tests on the first rotor profile were completed first on priority. Similar profiles have been designed for first rotor, stator and second rotor with minor changes in stagger angle and pitch. Hence, the same cascade blades, which were used to test the first rotor configuration, were deployed at pitch and stagger corresponding to the stator to test the equivalent configuration. The LOX stator configuration was tested at three different inlet flow angles and six outlet Mach numbers. Performance parameters such as profile loss, exit flow angle, flow velocities and surface Mach number distribution were evaluated. The test details and results of the LOX stator configuration are elaborated in this report

CASCADE TESTS ON CE 20 LOX TURBINE NOZZLE BLADE PROFILE

LPSC, ISRO, as part of their CE 20 cryogenic engine development program, desired to get the liquid oxygen [LOX] turbo pump turbine profiles tested at the NAL Transonic Cascade Tunnel to obtain basic aerodynamic performance data. The first rotor, stator and the second rotor profiles of the LOX turbine have already been tested in TCT and the results have been disseminated. The present task involves the aerodynamic evaluation of the nozzle profile of LOX turbine, in TCT. This profile was tested at five different inlet flow angles and eight outlet Mach numbers, covering the design and several off design conditions. Aerodynamic performance parameters such as profile loss, exit flow angle, flow velocities and surface Mach number distribution were evaluated. The effect of incidence on these performance parameters were also studied in detail. Oil flow visualization study was done at the design condition to study the flow behavior on the blade surfaces

CASCADE TESTS ON CE 20 LOX TURBINE SECOND ROTOR BLADE PROFILE

LPSC, ISRO as part of their CE 20 cryogenic engine development program, desired to get the liquid oxygen [LOX] turbo pump turbine profiles tested in the NAL Transonic Cascade Tunnel to obtain basic aerodynamic performance data. The LOX turbine consists of a nozzle, first rotor, stator and second rotor. Similar profiles have been designed for the first rotor, stator and second rotor with minor changes in stagger angle and pitch. Hence, the same cascade blades were used to test the first rotor, stator and second rotor configurations at respective pitch, stagger and flow conditions. The tests on the first rotor and stator configuration have been completed earlier. The test details and results of the LOX second rotor configuration are presented in this report. The LOX second rotor configuration was tested at three different inlet flow angles and six outlets Mach numbers. Performance parameters such as profile loss, exit flow angle, flow velocities and surface Mach number distribution were evaluated

EXPERIMENTAL INVESTIGATIONS ON SEMI-CRYO ROCKET ENGINE TURBO PUMP TURBINE PROFILES PART II - ROTOR PROFILE

LPSC, ISRO, as part of their semi-cryogenic rocket engine development program for the reusable launch vehicle, desired to get its turbo pump turbine profiles tested in the NAL Transonic Cascade Tunnel (TCT), to obtain basic aerodynamic performance data. The mean section profiles of the nozzle vane and rotor blade of this turbo pump turbine were tested in TCT. The test details and results of the rotor blade profile are discussed in this report viz a viz Part II and the details pertaining to the nozzle vane profile are furnished in Part I. The aerodynamic performance parameters such as profile loss, flow deflection, flow velocities and surface Mach number distribution were evaluated during the cascade tests. Both the profiles were tested over a range of inlet flow angles and outlet Mach numbers covering their respective design and off-design conditions. Detailed post processing of the test results were carried out and key aspects such as the effect of incidence and effect of Mach number on the performance were studied. An oil flow visualization study was also conducted on these profiles at their respective design condition, to ascertain the flow pattern over the blade surfaces

Aerodynamic Characteristics Of Elliptic Bodies At M = 2.0 And 3.0

Aerodynamic Data Is Generated On A Number Of Circular And Elliptic Bodies At Mach Numbers Of 2 .0 And 3.0 At Unit Reynolds Number Of 27 And 32 Million Per Meter Respectively. The Incidence Range Covered Was From -2 .0 To 10 Degree. The Bodies Tested Had Ellipticity Ratios Of 1 .0, 1 .5, 2 .0, 3 .0 Nose Fineness Ratios Of 1 .0, 2 .0, 3 .0 And Aft Body Ratios Of 5.0,6.0 And 7.0. Results Show Considerable Effect Of Ellipticity On The Normal Force Pitching Moment And Drag Coefficients.For A Body Of Constant Ellipticity Ratio, Increase In Nose Fineness Ratio Shows An Increase In The Lift Coefficient And Decrease In The Drag Coefficient,While The Increase In Aft Body Fineness Ratio Beyond 6.0 Shows A Decrease In The Drag Coefficient At Higher Incidence.Increase In Mach Number Showed A Reduction In Lift/Drag Ratio At Higher Incidences. Experimental Data Obtained Compared Well With The Theoretical Data Estimated Using ESDU Data Sheets

The NAL transonic cascade tunnel

A versatile test facility has been developed by the National Aeronautical Laboratory, Bangalore, India, which has the capability of testing turbomachinery blade cascades at transonic /supersonic inlet Mach numbers. The tunnel has a large test section and is equipped with13; suitable instrumentation, control, and safety systems. A description of this facility is presented, along with details of the instrumentation and control systems, and typical tunnel characteristics and calibration results are also given. It is concluded that the versatility of the tunnel makes it well equipped to take up a wide range of research and developmental activities connected with transonic flows inside turbomachine blade passages

Effect of Axial Velocity Density Ratio on the Performance of a Controlled Diffusion Airfoil Compressor Cascade

Axial Velocity Density Ratio (AVDR) is an important parameter to check the two-dimensionality of cascade flows. It can have significant influence on the cascade performance and the secondary flow structure. In the present study, the effect of AVDR has been investigated on a highly loaded Controlled Diffusion airfoil compressor cascade. Detailed 3D Computational Fluid Dynamics (CFD) studies were carried out with the cascade at five different AVDRs. Key aerodynamic performance parameters and flow structure through the cascade were analyzed in detail. CFD results of one AVDR were validated with the experimental cascade test data and were seen to be in good agreement. Loss characteristics of the cascade varied significantly with change in AVDR. Increase in AVDR postponed the point of separation on the suction surface, produced thinner boundary layers and caused substantial drop in the pressure loss coefficient. Strong end wall vortices were noticed at AVDR of 1.177. At higher AVDRs, the flow was well guided even close to the end wall and the secondary flows diminished. The loading initially improved with increase in AVDR. Beyond a certain limit, further increase in AVDR offered no improvements to the loading but rather resulted in drop in diffusion and deviation