Coolant Gas Injection on a Blunt-Nosed Re-Entry Vehicle

2013 ◽  
Author(s):  
Jeswin Joseph ◽  
S. R. Shine
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Thermal Protection ◽  
Thermal Protection System ◽  
Protection System ◽  
Aerodynamic Heating ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Mach Numbers

Very high thermal loads are expected in re-entry vehicles traveling at hypersonic Mach numbers due to severe aerodynamic heating. In the present study, numerical investigations are carried out to analyze the use of film cooling technology for a fully reusable and active thermal protection system of the re-entry vehicle. Simulations are done to examine the fundamental flow phenomenon and the performance of blunt body film cooling in hypersonic flows. Simulations are conducted for a blunt -nosed spacecraft flying at Mach numbers varying from 4 to 8 and 40 deg angle of attack. Film cooling holes are provided on the bottom of the blunt-nosed body. Standard values at an altitude of 30 km are used as in flow boundary conditions. The dependency of blowing ratios, stream-wise injection angle and inlet Mach number on the film cooling effectiveness are investigated. It is observed that the film cooling effectiveness reduces with increase in coolant injection angle. The film cooling performance is found to be decreasing with increase in Mach number. The results could provide useful inputs for optimization of an active thermal protection system of re-entry vehicles.

scholarly journals Adiabatic Wall Effectiveness Measurements of Film-Cooling Holes With Expanded Exits

1997 ◽  
Author(s):  
M. Gritsch ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Thermal Protection ◽  
Lateral Spreading ◽  
Cylindrical Hole ◽  
Adiabatic Wall ◽  
Camera System ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole

This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser shaped exit portion (i.e. a fanshaped and a laidback fanshaped hole). The flow conditions considered are the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the blowing ratio (up to 2). The coolant-to-mainflow temperature ratio is kept constant at 0.54. The measurements are performed by means of an infrared camera system which provides a two-dimensional distribution of the film-cooling effectiveness in the nearfield of the cooling hole down to x/D = 10. As compared to the cylindrical hole, both expanded holes show significantly improved thermal protection of the surface downstream of the ejection location, particularly at high blowing ratios. The laidback fanshaped hole provides a better lateral spreading of the ejected coolant than the fanshaped hole which leads to higher laterally averaged film-cooling effectiveness. Coolant passage crossflow Mach number and orientation strongly affect the flowfield of the jet being ejected from the hole and, therefore, have an important impact on film-cooling performance.

Effectiveness of Normal and Angled Slot Cooling

2004 ◽  
Author(s):  
A. C. Smith ◽  
J. H. Hatchett ◽  
A. C. Nix ◽  
W. F. Ng ◽  
K. A. Thole ◽  
...  
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Fluid Dynamic ◽  
Current Experiment ◽  
Cfd Simulations ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Mach Numbers ◽  
Lift Off

An experimental and numerical investigation was conducted to determine the film cooling effectiveness of a normal slot and angled slot under realistic engine Mach number conditions. Freestream Mach numbers of 0.65 and 1.3 were tested. For the normal slot, hot gas ingestion into the slot was observed at low blowing ratios (M < 0.25). At high blowing ratios (M > 0.6) the cooling film was observed to “lift off” from the surface. For the 30° angled slot, the data was found to collapse using the blowing ratio as a scaling parameter. Results from the current experiment were compared with the subsonic data previously published. For the angle slot, at supersonic freestream Mach number, the current experiment shows that at the same x/Ms, the film-cooling effectiveness increases by as much as 25% as compared to the subsonic case. The results of the experiment also show that at the same x/Ms, the film cooling effectiveness of the angle slot is considerably higher than the normal slot, at both subsonic and supersonic Mach numbers. The flow physics for the slot tests considered here are also described with computational fluid dynamic (CFD) simulations in the subsonic and supersonic regimes.

Adiabatic Wall Effectiveness Measurements of Film-Cooling Holes With Expanded Exits

1998 ◽  
Vol 120 (3) ◽  
pp. 549-556 ◽  
Author(s):  
M. Gritsch ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Thermal Protection ◽  
Lateral Spreading ◽  
Cylindrical Hole ◽  
Camera System ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Shaped Hole

This paper presents detailed measurements of the film-cooling effectiveness for three single, scaled-up film-cooling hole geometries. The hole geometries investigated include a cylindrical hole and two holes with a diffuser-shaped exit portion (i.e., a fan-shaped and a laid-back fan-shaped hole). The flow conditions considered are the crossflow Mach number at the hole entrance side (up to 0.6), the crossflow Mach number at the hole exit side (up to 1.2), and the blowing ratio (up to 2). The coolant-to-mainflow temperature ratio is kept constant at 0.54. The measurements are performed by means of an infrared camera system, which provides a two-dimensional distribution of the film-cooling effectiveness in the near field of the cooling hole down to x/D = 10. As compared to the cylindrical hole, both expanded holes show significantly improved thermal protection of the surface downstream of the ejection location, particularly at high blowing ratios. The laidback fan-shaped hole provides a better lateral spreading of the ejected coolant than the fan-shaped hole, which leads to higher laterally averaged film-cooling effectiveness. Coolant passage cross-flow Mach number and orientation strongly affect the flowfield of the jet being ejected from the hole and, therefore, have an important impact on film-cooling performance.

Film Cooling Effect on the Nosetip in Hypersonic Flow

2013 ◽  
Vol 432 ◽  
pp. 368-372
Author(s):  
Tao Nie ◽  
Wei Qiang Liu
Keyword(s):  
Film Cooling ◽  
Free Stream ◽  
Thermal Protection ◽  
Aerodynamic Force ◽  
Aerodynamic Resistance ◽  
Thermal Protection System ◽  
Protection System ◽  
Cooling Effect ◽  
Recirculation Region ◽  
Aerodynamic Heating

The cooling effect of a film cooling thermal protection system is investigated, and the flow field, aerodynamic force and surface temperature distribution are obtained. The numerical method is validated by experiment with no film cooling. The physical mechanism of the reduction of temperature is analyzed. The jet interacts with the free stream to form a thermal insulating layer on the wall, which enables the free stream to flow outside the wall rather than interact with the surface to produce aerodynamic heating. In addition, the film cooling flow form a cool recirculation region, which reduces the temperature on the surface. Analysis of the numerical simulation results shows that this kind of thermal protection system has an excellent cooling effect on the surface of the nose-tip. With the outlet speed increasing, the cooling efficiency is improved and the aerodynamic resistance is changeless. When outlet speed and total pressure are invariable, the cooling effect of air and nitrogen is the same, but the cooling effect of helium is better.

Numerical Investigation of Effects of Blockage, Inclination Angle, and Hole-Size on Film Cooling Effectiveness at Concave Surface

2021 ◽  
Vol 143 (2) ◽  
Author(s):  
Fu-qiang Wang ◽  
Jian Pu ◽  
Jian-hua Wang ◽  
Wei-dong Xia
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Numerical Study ◽  
Density Ratio ◽  
Concave Surface ◽  
Blockage Ratio ◽  
Coverage Area ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle

Abstract Film-hole can be often blocked by thermal-barrier coatings (TBCs) spraying, resulting in the variations of aerodynamic and thermal performances of film cooling. In this study, a numerical study of the blockage effect on the film cooling effectiveness of inclined cylindrical-holes was carried out on a concave surface to simulate the airfoil pressure side. Three typical blowing ratios (BRs) of 0.5, 1.0, and 1.5 were chosen at an engine-similar density ratio (DR) of 2.0. Two common inclination angles of 30 deg and 45 deg were designed. The blockage ratios were adjusted from 0 to 20%. The results indicated the blockage could enhance the penetration of film cooling flow to the mainstream. Thus, the averaged effectiveness and coolant coverage area were reduced. Moreover, the pressure loss inside of the hole was increased. With the increase of BR, the decrement of film cooling effectiveness caused by blockage rapidly increased. At BR = 1.5, the decrement could be acquired up to 70% for a blockage ratio of 20%. The decrement of film cooling effectiveness caused by blockage was nearly nonsensitive to the injection angle; however, the larger angle could generate the higher increment of pressure loss caused by blockage. A new design method for the couple scheme of film cooling and TBC was proposed, i.e., increasing the inlet diameter according to the blockage ratio before TBC spraying. In comparison with the original unblocked-hole, the enlarged blocked-hole not only kept the nearly same area-averaged effectiveness but also reduced slightly the pressure loss inside of the hole. Unfortunately, application of enlarged blocked-hole at large BR could lead to a more obvious reduction of effectiveness near hole-exit, in comparison with the original common-hole.

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.

Injection Angle Influence of Secondary Holes on the Enhancement of Film Cooling Effectiveness With Horn-Shaped or Cylindrical Primary Holes

2017 ◽  
Author(s):  
Rui Zhu ◽  
Gongnan Xie ◽  
Terrence W. Simon
Keyword(s):  
Film Cooling ◽  
Numerical Study ◽  
Cylindrical Hole ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
Inclination Angles ◽  
Cooling Hole ◽  
Cylindrical Holes

Secondary holes to a main film cooling hole are used to improve film cooling performance by creating anti-kidney vortices. The effects of injection angle of the secondary holes on both film cooling effectiveness and surrounding thermal and flow fields are investigated in this numerical study. Two kinds of primary hole shapes are adopted. One is a cylindrical hole, the other is a horn-shaped hole which is designed from a cylindrical hole by expanding the hole in the transverse direction to double the hole size at the exit. Two smaller cylindrical holes, the secondary holes, are located symmetrically about the centerline and downstream of the primary hole. Three compound injection angles (α = 30°, 45° and 60°, β = 30°) of the secondary holes are analyzed while the injection angle of the primary hole is kept at 45°. Cases with various blowing ratios are computed. It is shown from the simulation that cooling effectiveness of secondary holes with a horn-shaped primary hole is better than that with a cylindrical primary hole, especially at high blowing ratios. With a cylindrical primary hole, increasing inclination angle of the secondary holes provides better cooling effectiveness because the anti-kidney vortices created by shallow secondary holes cannot counteract the kidney vortex pairs adequately, enhancing mixing of main flow and coolant. For secondary holes with a horn-shaped primary hole, large secondary hole inclination angles provide better cooling performance at low blowing ratios; but, at high blowing ratios, secondary holes with small inclination angles are more effective, as the film coverage becomes wider in the downstream area.

Systematic Study of Film Cooling With a Three-Dimensional Calculation Procedure

1986 ◽  
Vol 108 (1) ◽  
pp. 124-130 ◽  
Author(s):  
A. O. Demuren ◽  
W. Rodi ◽  
B. Scho¨nung
Keyword(s):  
Systematic Study ◽  
Film Cooling ◽  
Three Dimensional ◽  
Heat Fluxes ◽  
Test Cases ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Injection Angle ◽  
The Individual ◽  
Volume Method

The present paper describes three-dimensional calculations of film cooling by injection from a single row of holes. A systematic study of the influence of different parameters on the cooling effectiveness has been carried out. Twenty-seven test cases have been calculated, varying the injection angle (α = 10/45/90 deg), the relative spacing (s/D = 1.5/3/5) and the blowing rate (M = 0.5/1/2) for the same mainstream conditions. The governing three-dimensional equations are solved by a finite volume method. The turbulent stresses and heat fluxes are obtained from a k–ε model modified to account for nonisotropic eddy viscosities and diffusivities. Examples of predicted velocity and temperature distributions are presented and compared with available experimental data. For all the test cases, the laterally averaged cooling effectiveness is given. On the whole, the agreement with experiments is fairly good, even though there are discrepancies about details in some of the cases. The influence of the individual parameters on the film cooling effectiveness is predicted correctly in all cases. This influence is discussed in some detail and the parameter combination with the best overall cooling performance is identified.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

scholarly journals Thermomechanical Performance of Bio-Inspired Corrugated-Core Sandwich Structure for a Thermal Protection System Panel

2019 ◽  
Vol 9 (24) ◽  
pp. 5541 ◽  
Author(s):  
Vinh Tung Le ◽  
Nam Seo Goo
Keyword(s):  
Sandwich Structure ◽  
Thermal Protection ◽  
Sandwich Structures ◽  
High Strength ◽  
Thermal Protection System ◽  
Boundary Layer Transition ◽  
Protection System ◽  
Aerodynamic Heating ◽  
Dense Structure ◽  
Skin Structure

A skin structure for thermal protection is one of the most interesting components that needs to be considered in the design of a hypersonic vehicle. The thermal protection structure, if a dense structure is used, is heavy and has a large heat conduction path. Thus, a lightweight, high strength structure is preferable. Currently, for designing a lightweight structure with high strength, natural materials are of great interest for achieving low density, high strength, and toughness. This paper presents bio-inspired lightweight structures that ensure high strength for a thermal protection system (TPS). A sinusoidal shape inspired by the microstructure of the dactyl club of Odontodactylus scyllarus, known as the peacock mantis shrimp, is presented with two different geometries, a unidirectionally corrugated core sandwich structure (UCS) and a bidirectionally corrugated core sandwich structure (BCS). Thermomechanical analysis of the two corrugated core structures is performed under simulated aerodynamic heating, and the total deflection and thermal stress are presented. The maximum deflection of the present sandwich structure throughout a mission flight was 1.74 mm for the UCS and 2.04 mm for the BCS. Compared with the dense structure used for the skin structure of the TPS, the bio-inspired corrugated core sandwich structures achieved about a 65% weight reduction, while the deflections still satisfied the limits for delaying the hypersonic boundary layer transition. Moreover, we first fabricated the BCS to test the thermomechanical behaviors under a thermal load. Finally, we examined the influence of the core thickness, face-sheet thickness, and emittance in the simulation model to identify appropriate structural parameters in the TPS optimization. The present corrugated core sandwich structures could be employed as a skin structure for metallic TPS panels instead of the honeycomb sandwich structure.

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