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|Title: ||Investigations On Film Cooling At Hypersonic Mach Number Using Forward Facing Injection From Micro-Jet Array|
|Authors: ||Sriram, R|
|Advisors: ||Jagadeesh, G|
|Keywords: ||Hypersonic Aerodynamics|
Hypersonic Mach Number
Hypersonic Mach Numbers
|Submitted Date: ||1-Aug-2008|
|Series/Report no.: ||G22441|
|Abstract: ||A body in a hypersonic flow field will experience very high heating especially during re-entry. Conventionally this problem is tackled to some extent by the use of large angle
blunt cones. At the cost of increased drag, the heat transfer rate is lower over most parts of the blunt body, except in a region around the stagnation point. Thus even with blunt cones, management of heat transfer rates and drag on bodies at hypersonic speeds
continues to be an interesting research area. Various thermal protection systems have
been proposed in the past, like heat sink cooling, ablation cooling and aerospikes. The
ablative cooling system becomes extremely costly when reusability is the major concern.
Also the shape change due to ablation can lead to issues with the vehicle control. The
aerospikes themselves may become hot and ablate at hypersonic speeds. Hence an
alternate form of cooling system is necessary for hypersonic flows, which is more
feasible, cost effective and efficient than the conventional cooling systems.
Injection of a mass of cold fluid into the boundary layer through the surface is one
of the potential cooling techniques in the hypersonic flight corridors. These kinds of
thermal protection systems are called mass transfer cooling systems. The injection of the mass may be through discrete slots or through a porous media. When the coolant is
injected through a porous media over the entire surface, the coolant comes out as a
continuous mass. Such a cooling system is also referred as “transpiration cooling
system”. When the fluid is injected through discrete slots, the system is called as “film
cooling system”. In either case, the coolant absorbs the incoming heat through its rise in
enthalpy and thus modifies the boundary layer characteristics in such a way that the heat flow rate to the surface is less. Injection of a forward facing jet (opposite to the freestream direction) from the stagnation point of a blunt body can be used for mitigating both the aerodynamic drag and heat transfer rates at hypersonic Mach numbers. If the jet has enough momentum it can push the bow shock forward, resulting in reduced drag. This will also reduce heat transfer rate over most part of the body except around the jet re-attachment region. A reattachment shock impinging on the blunt body invariably increases the local heat flux. At lower momentum fluxes the forward facing jet cannot push the bow shock ahead of the blunt body and spreads easily over the boundary layer, resulting in reduced heat transfer rates. While the film cooling performance improves with mass flow rate of the jet, higher momentum flow rates can lead to a stronger reattachment leading to higher heat transfer rate at the reattachment zone. If we are able to reduce the momentum flux of the coolant for the same mass flow rate, the gas coming out can easily spread over the boundary layer and it is possible to improve the film cooling performance.
In all the reported literature, the mass flow rate and the momentum flux are not
varied independently. This means, if the mass flow rate is increased, there is a
corresponding increase in the momentum flux. This is because the injection (from a
particular orifice and for a particular coolant gas) is controlled only by the total pressure of injection and free stream conditions. The present investigation is mainly aimed at demonstrating the effect of reduction in momentum of the coolant (injected opposing a hypersonic freestream from the stagnation point of a blunt cone), keeping the mass flow rate the same, on the film cooling performance. This is achieved by splitting a single jet into a number of smaller jets of same injection area (for same injection total pressure and same free stream conditions). To the best of our knowledge there is no report on the use
of forward facing micro-jet array for film cooling at hypersonic Mach numbers. In this
backdrop the main objectives of the present study are:
• To experimentally demonstrate the effect of splitting a single jet into an array of closely spaced smaller micro-jets of same effective area of injection (injected opposite to a hypersonic freestream from the stagnation zone), on the reduction in surface heat transfer rates on a large angle blunt cone.
· Identifying various parameters that affect the flow phenomenon and doing a systematic investigation of the effect of the different parameters on the surface heat transfer rates and drag.
Experimental investigations are carried out in the IISc hypersonic shock tunnel on
the film cooling effectiveness. Coolant gas (nitrogen and helium) is injected opposing
hypersonic freestream as a single jet (diameter 2 mm and 0.9 mm), and as an array of iv micro jets (diameter 300 micron each) of same effective area (corresponding to the
respective single jet). The coolant gas is injected from the stagnation zone of a blunt cone model (58o apex angle and nose radius of 35 mm). Experiments are performed at a flow freestream Mach number of 5.9 at 0o angle of attack, with a stagnation enthalpy of 1.84 MJ/Kg, with and without injections. The ratios of the jet stagnation pressure to the pitot pressure (stagnation pressure ratio) used in the present study are 1.2 and 1.45. Surface convective heat transfer measurements using platinum thin film sensors, time resolved schlieren flow visualization and aerodynamic drag measurements using accelerometer force balance are used as flow diagnostics in the present study. The theoretical stagnation
point heat transfer rate without injection for the given freestream conditions for the test model is 79 W/cm2 and the corresponding aerodynamic drag from Newtonian theory is
143 N. The measured drag value without injection (125 N) shows a reasonable match
with theory. As the injection is from stagnation zone it is not possible to measure the surface heat transfer rates at the stagnation point. The sensors thus are placed from the nearest possible location from the stagnation point (from 16 mm from stagnation point on the surface). The sensors near the stagnation point measures a heat transfer rate of 65 W/cm2 on an average without any injection. Some of the important conclusions from the study are:
• Up to 40% reduction in surface heat transfer rate has been measured near the
stagnation point with the array of micro jets, nitrogen being the coolant, while the
corresponding reduction was up to 30% for helium injection. Considering the single jet injection, near the stagnation point there is either no reduction in heat transfer rate or a slight increase up to 10%.
· Far away from stagnation point the reduction in heat transfer with array of micro-jets is only slightly higher than corresponding single jet for the same pressure ratio. Thus the cooling performance of the array of closely spaced micro jets is
better than the corresponding single jet almost over the entire surface.
• The time resolved flow visualization studies show no major change in the shock
standoff distance with the low momentum gas injection, indicating no major changes in other aerodynamic aspects such as drag.
· The drag measurements also indicate that there is virtually no change in the overall aerodynamic drag with gas injection from the micro-orifice array.
· The spreading of the jets injected from the closely spaced micro-orifice array over
the surface is also seen in the visualization, indicating the absence of a region of strong reattachment.
· The reduction in momentum flux of the injected mass due to the interaction
between individual jets in the case of closely spaced micro-jet array appears to be
the main reason for better performance when compared to a single jet.
The thesis is organized in six chapters. The importance of film cooling at hypersonic speeds and the objectives of the investigation are concisely presented in
Chapter 1. From the knowledge of the flow field with counter-flow injection obtained
from the literature, the important variables governing the flow phenomena are organized
as non-dimensional parameters using dimensional analysis in Chapter 2. The description of the shock tunnel facility, diagnostics and the test model used in the present study is given in Chapter 3. Chapter 4 describes the results of drag measurements and flow visualization studies. The heat transfer measurements and the observed trends in heat transfer rates with and without coolant injection are then discussed in detail in Chapter 5. Based on the obtained results the possible physical picture of the flow field is discussed
in Chapter 6, followed by the important conclusions of the investigation.|
|Appears in Collections:||Aerospace Engineering (aero)|
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