Effects of shock waves interaction on hydrocarbon fueled supersonic film cooling with combustion

2021 ◽  
pp. 106693
Author(s):  
Jingying Zuo ◽  
Silong Zhang ◽  
Jiang Qin ◽  
Wen Bao ◽  
Naigang Cui ◽  
...  
Keyword(s):  
Shock Waves ◽  
Film Cooling ◽  
Waves Interaction

ON SIMULATING ACOUSTIC-SHOCK WAVES INTERACTION

2001 ◽  
Vol 9 (9) ◽  
pp. 973-984
Author(s):  
KHALID HOSNY ◽  
ISMAIL ISMAIL ◽  
AWATEF NAMED
Keyword(s):  
Shock Waves ◽  
Waves Interaction

Shock Wave - Film Cooling Interactions in Transonic Flows

2001 ◽  
Author(s):  
P. M. Ligrani ◽  
C. Saumweber ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Waves ◽  
Film Cooling ◽  
Cross Flow ◽  
Injection Velocity ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Density Ratios ◽  
Film Cooling Holes

Interactions between shock waves and film cooling are described as they affect magnitudes of local and spanwise-averaged adiabatic film cooling effectiveness distributions. A row of three cylindrical holes is employed. Spanwise spacing of holes is 4 diameters, and inclination angle is 30 degrees. Freestream Mach numbers of 0.8 and 1.10–1.12 are used, with coolant to freestream density ratios of 1.5–1.6. Shadowgraph images show different shock structures as the blowing ratio is changed, and as the condition employed for injection of film into the cooling holes is altered. Investigated are film plenum conditions, as well as perpendicular film injection cross-flow Mach numbers of 0.15, 0.3, and 0.6. Dramatic changes to local and spanwise-averaged adiabatic film effectiveness distributions are then observed as different shock wave structures develop in the immediate vicinity of the film-cooling holes. Variations are especially evident as the data obtained with a supersonic Mach number are compared to the data obtained with a freestream Mach number of 0.8. Local and spanwise-averaged effectiveness magnitudes are generally higher when shock waves are present when a film plenum condition (with zero cross-flow Mach number) is utilized. Effectiveness values measured with a supersonic approaching freestream and shock waves then decrease as the injection cross-flow Mach number increases. Such changes are due to altered flow separation regions in film holes, different injection velocity distributions at hole exits, and alterations of static pressures at film hole exits produced by different types of shock wave events.

Head-on collision of two spherical shock waves. Interaction of laser sparks in a gas

1991 ◽  
Vol 25 (5) ◽  
pp. 761-765
Author(s):  
V. A. Andrushchenko ◽  
S. Yu. Efimov ◽  
L. A. Chudov
Keyword(s):  
Shock Waves ◽  
Waves Interaction ◽  
Spherical Shock

scholarly journals ON SIMULATING ACOUSTIC-SHOCK WAVES INTERACTION

2001 ◽  
Vol 9 (ASAT Conference, 8-10 May 2001) ◽  
pp. 1-12
Author(s):  
KHALID HOSNY ◽  
ISMAIL ISMAIL ◽  
AWATEF NAMED
Keyword(s):  
Shock Waves ◽  
Waves Interaction

Computational Study of the Effects of Shock Waves on Film Cooling Effectiveness

2009 ◽  
Author(s):  
Chad X.-Z. Zhang ◽  
Ibrahim G. Hassan
Keyword(s):  
Shock Waves ◽  
Flat Plate ◽  
Film Cooling ◽  
Computational Study ◽  
Strong Shock ◽  
Boundary Layer Separation ◽  
Airfoil Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Transonic Airfoil

The performance of a louver cooling scheme on a transonic airfoil has been studied numerically in this paper. Film cooling holes are located near the passage throat. The Mach number at the location of the jet exit is close to unity. A comparison of film cooling effectiveness between numerical prediction and experimental data for a circular hole shows that the numerical procedures are adequate. In addition to the shock wave effects and compressibility, curvature effect was also studied by comparing cooling effectiveness on the airfoil surface with that on a flat plate. Substantially higher cooling effectiveness for the louver cooling scheme on the airfoil was predicted at blowing ratios below 1 in comparison to other cooling configurations. At higher blowing ratios than 2 the advantages of the louver cooling scheme becomes less obvious. It was also found that for the same cooling configuration the cooling effectiveness on the transonic airfoil is slightly higher than that on a flat plate at moderately low blowing ratios below 1. At high blowing ratios above 2 when the oblique shock becomes detached from the leading edge of the hole exits, dramatic reduction in cooling effectiveness occurs as a result of boundary layer separation due to the strong shock waves. A coolant-blockage and shaped-wedge analogy was proposed and found to be able to qualitatively explain this phenomenon satisfactorily.

Simulation of Air/Mist Cooling Among Shock Waves and Passing Wakes Interactions in a Transonic Gas Turbine Stage

2021 ◽  
Author(s):  
Ting Wang ◽  
Ramy Abdelmaksoud
Keyword(s):  
Shock Waves ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Water Droplets ◽  
Discrete Phase Model ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Enhancement ◽  
Mist Cooling

Abstract This paper presents a 2-D numerical investigation of the effect of interactions of moving wakes and shock waves on mist cooling performance over airfoils in the first stator-rotor stage of a transonic gas turbine. The discrete phase model (DPM) is used to simulate and track the evaporation and movement of the tiny water droplets. Breakup and coalescence sub-models are used to simulate the interaction between the droplets themselves. A linear sliding mesh technique is used to study the transient stator-rotor interaction. The results show that the passing unsteady wakes caused by the blade rotation press the mist on the blade suction side flowing near the blade surface, providing more enhanced film cooling effectiveness. The weak oblique shock waves do not exert a significant effect on the air/mist cooling effectiveness. Injecting a 10% mist ratio noticeably improved the cooling enhancement by reducing the wall temperature values up to 200 K in some locations. Injecting the tiny water droplets does not cause a noticeable pressure loss compared to the air-only cooling case. Injecting mist doesn’t alter the effect of shocks.

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.

Influence of Shock Waves on Supersonic Film Cooling

2009 ◽  
Vol 46 (1) ◽  
pp. 67-73 ◽  
Author(s):  
Wei Peng ◽  
Pei-Xue Jiang
Keyword(s):  
Shock Waves ◽  
Film Cooling
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