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The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag reduction of the vehicle along with the minimized heat transfer rate. Using forward facing solid aerospike is the simplest way among the existing drag reduction methodologies for hypersonic blunt cone bodies. But the flow oscillations associated with this aerospike makes it difficult to implement. When analyzing this flow, it can be understood that this oscillating flow can be compared to conical cavity flow. Therefore in the spiked flows, it is decided to implement the technique used in reducing the flow oscillation of the cavities. Based on this method the shallow conical cavity flow generated by the aerospike fixed ahead of a 120o blunt cone body is fissured as multiple cavities by so many disks formed from 10o cone. Now the deep conical cavities had the length to mean depth ratio of unity; this suppresses the unnecessary oscillations of the shallow cavity. The total length of the telescopic aerospike is fixed as 100mm. And one another conical tip plain aerospike of same length is designed for comparing the telescopic spike’s performance at hypersonic flow Mach numbers of 5.75 and 7.9. A three component force balance system capable of measuring drag, lift and pitching moment is designed and mounted internally into the skirt of the model. Drag measurement is done for without spike, conical tip plain spiked and telescopic spiked blunt cone body. The three configurations are tested at different angles of attack from 0 to 10 degree with a step of 2. A discrete iterative deconvolution methodology is implemented in this research work for obtaining the clean drag history from the noisy drag accelerometer signal. The drag results showed the drag reduction when compared to the without spike blunt cone body. When comparing to the plain spiked, the telescopic spiked blunt cone body has lesser drag at higher angles of attack. Heat transfer measurements are done over the blunt cone surface using the Platinum thin film gauges formed over the Macor substrate. These results and the flow visualization give better understanding of the flow and the heat flux rate caused by the flow. The enhancement in the heat flux rate over the blunt cone surface is due to the shock interaction. And in recirculation region the heat flux rate is very much lesser when compared to without spike blunt cone body. It is observed that the shock interaction in the windward side is coming closer towards the nose of the blunt cone as the angle of attack increases and the oscillation of the oblique shock also decreases. Schlieren visualization showed that there is dispersion in the oblique shock, particularly in the leeward side. In the telescopic spike there are multiple shocks generated from each and every disk which coalesces together to form a single oblique shock. And the effect of the shock generated by the telescopic spike is stronger than the effect of the shock generated by the conical tip plain spike.

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.

In the flying field of space transportation domain, the increased efforts involving design and development of hypersonic flight for space missions is on toe to provide the optimum aerothermodynamic design data to satisfy mission requirements. Aerothermodynamics is the basis for designing and development of hypersonic space transportation flight vehicles such as X 51 a, and other programmes like planetary probes for Moon and Mars, and Earth re-entry vehicles such as SRE and space shuttle. It enables safe flying of aerospace vehicles, keeping other parameters optimum for structural and materials with thermal protection systems. In this context, the experimental investigations on hypersonic waverider are carried out at design Mach 6. The hypersonic waverider has high lift to drag ratio at design Mach number even at zero degree angle of incidence, and this seems to be one of the special characteristics for its shape at hypersonic flight regime. The heat transfer rates are measured using 30 thin film platinum gauges sputtered on a Macor material that are embedded on the test model. The waverider has 16 sensors on top surface and 14 on bottom surface of a model. The surface temperature history is directly converted to heat transfer rates. The heat transfer data are measured for design (Mach 6) and off-design Mach numbers (8) in the hypersonic shock tunnel, HST2. The results are obtained at stagnation enthalpy of ~ 2 MJ/kg, and Reynolds number range from 0.578 x 106 m-1 to 1.461 x 106 m-1. In addition, flow visualization is carried out by using Schlieren technique to obtain the shock structures and flow evolution around the Waverider. Some preliminary computational analyses are conducted using FLUENT 6.3 and HiFUN, which gave quantitative results. Experimentally measured surface heat flux data are compared with the computed one and both the data agree well. These detailed results are presented in the thesis.

One of the major technological innovations that are necessary for faster and cheaper access-to-space will be the commercial realization of supersonic combustion jet engines (SCRAMJET). The establishment of the flow through the inlet is one the prime requirement for the success of a SCRAMJET engine. The flow through a SCRAMJET inlet is dominated by inviscid /viscous coupling, transition, shock-shock interaction, shock boundary layer interaction, blunt leading edge effects and flow profile effects. Although the literature is exhaustive on various aspects of flow features associated with SCRAMJET engines, very little is known on the fundamental gasdynamic features dictating the flow establishment in the SCRAMJET inlet. On one hand we need the reduction of flight Mach number to manageable supersonic values inside the SCRAMJET combustor, but on the other hand we have to achieve this with minimum total pressure loss. Hence the dynamics of ramp/cowl shock interaction process ahead of the inlet has a direct bearing on the quality and type of flow inside the SCRAMJET engine. There is virtually no data base in the open literature focusing specifically on the cowl/ramp shock interactions at hypersonic Mach numbers. Hence in this backdrop, the main aim of the present investigation is to systematically understand the ramp/cowl shock interaction processes in front of a generic inlet model. Since we are primarily concerned with the shock interaction process ahead of the cowl all the investigations are carried out without any fuel injection. Variable geometry is necessary if we want to operate the inlet for a wide range of Mach numbers in actual flight. The investigation mainly comprises of three variable geometry configurations; namely, variation of contraction ratios at 00 cowl (CR 8.4, 5.0 and 4.3), variation of cowl length for a given chamber height (four lengths of cowls at 10 mm chamber height) and variation of cowl angle (three angles cowl each for two chamber heights). The change in cowl configuration results in different ramp/cowl shock interaction processes affecting the performance of the inlet. Experiments are performed in IISc hypersonic shock tunnel HST 2 (test time ~ 1 ms) at two nominal Mach numbers 8.0 and 5.74 for design and off-design testing conditions. Exhaustive numerical simulations are also performed to compliment the experiments. Further the effect of concentrated energy deposition on forebody /cowl shock interactions has also been investigated. A 2D, planar, single ramp scramjet inlet model has been designed and fabricated along with various cowl geometries and tested in a hypersonic shock tunnel to characterize the forebody/cowl shock interaction process for different inlet configurations. Further a DC plasma power unit and a plasma torch have been designed, developed and fabricated to serve as energy source for conducting flow-alteration experiments in the inlet model. The V-I characteristics of the plasma torch is studied and an estimation of plasma temperature is also performed as a part of characterizing the plasma flame. Initial standardization experiments of blunt body flow field alteration using the plasma torch and hence its drag reduction, are performed to check the torch’s suitability to be used as a flow-altering device in a shock tunnel. The plasma torch is integrated successfully with the inlet model in a shock tunnel to perform experiments with plasma jet as the energy source. The above experiments are first of its kind to be conducted in a shock tunnel. They are performed at various pressure ratios and supply currents. Time resolved schlieren flow visualization using Phantom 7.1 (Ms Vision Research USA) high speed camera, surface static pressure measurements inside a generic inlet using miniature kulite transducer and surface convective heat transfer rate measurements inside a generic inlet using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study. Some of the important conclusions from the study are: • Experiments performed at different contraction ratios show different shock patterns. At CR 8.4, the SOL condition is satisfied, but the flow gets choked due to over contraction and flow through inlet is not established. For CR 5.0, formation of a small Mach stem before the chamber is observed with the reflection point on the cowl and the weak reflected shock entering inside the chamber. The Mach stem grows with time. For CR 4.3, the forebody/cowl shock interference created an Edney’s Type II shock interaction pattern. However, at off-design conditions, for CR 5 the shock reflection is regular and at CR 4.3, the Edney’s Type II pattern lasts for a short time. • For all lengths of cowl tested, 131mm and 141mm showed Edney’s Type II shock interference where as 151mm showed Edney’s Type I pattern at design condition. In all cases the flow is choked for high contraction ratio. At off-design condition these shock patterns do not last for the entire test time but rather it becomes a lambda pattern with the normal shock before the inlet. • For inlet configurations with cowl angle other than 00, the flow is found to be established for all cases at designed condition and except for 100 cowl at off-design condition. • For CR 8.4 the peak value of pressure (~1.7x104 Pa) occurs at a location of 151mm, where as for CR 5.0 and 4.3 they occur at 188mm and 206mm having values ~1.6x104 Pa and ~1.4x104 Pa respectively. These locations indicate the likely locations of shock impingements inside the chamber. • For cowl angle of 00 for a 10 mm chamber the maximum pressure value recorded is ~1.7x104 Pa whereas for 100 and 200 cowl it is ~1.1x104 Pa and 1.2 x104 Pa respectively. This is because in the first case the inlet is choked because of over contraction whereas in the other two cases the CR is less and flow is established inside the inlet. • The average heat transfer rates of last four heat transfer gauges (180 mm, 190 mm, 200 mm and 210 mm from the forebody tip) for all lengths of cowls tested are found to be almost same (~ 20 W/cm2). This is because the flow is choked in all these cases. The numerical simulation also shows uniform distribution here, consistent with the experimental findings. • The locations of heat transfer peaks for 100 cowl at design condition can be observed to be occurring at 170 mm and 200 mm from the forebody tip having values ~44 W/cm2 and ~39 W/cm2 respectively. For a 200 cowl they seem to be occurring at 170 mm and 180 mm from the forebody tip having values ~50 W/cm2 and ~30 W/cm2. These locations indicate the likely locations of shock impingements inside the chamber. With the evolution of concept of upstream fuel injection in recent times these may the most appropriate locations for fuel injection. • At higher jet pressure ratios the plasma jet/ramp shock interaction results in a lambda shock pattern with the triple point forming vertically above the cowl level. This means the normal shock stands in front of the inlet making a part of the flow entering the inlet subsonic. The reflected shock from the triple point also separates the ramp boundary layer. • At lower jet pressure ratios the triple point is formed below the cowl level and the flow entering inside the inlet is supersonic. The reflected shock interacts with the cowl shock and a weak separation shock is seen. • Experiments are performed with concentrated DC electric discharge as energy source. Even though the amount of energy dumped here is less than 0.25% of the total energy it creates a perceptible disturbance in the flow. • Experiments are also performed to see the effect of electric discharge as energy source on height of Mach stem for a given inlet configuration. Deposition of energy in the present location does not seem to alter the Mach stem height. However more experiments need to be performed by varying the energy location to see its effect. Non-intrusive energy sources like microwave and lasers can be thought of as options for depositing energy to study its effect on Mach stem height. Since they provide more flexibility on varying the location of energy the optimum location of energy can be found out for highest reduction of Mach stem height.

Aerodynamic cavities are common features on hypersonic vehicles which are caused in both large and small scale features like surface defects, pitting, gap in joints etc. In the hypersonic regime, the presence of such cavities alters the flow phenomenon considerably and heating rates adjacent to the discontinuities can be greatly enhanced due to the diversion of flow. Since the 1960s, a great deal of theoretical and experimental research has been carried out on cavity flow physics and heating. However, most of the studies have been done to characterize the effect downstream and within the cavity. In the present study, a series of were carried out in the shock tunnel to investigate the heating characteristics, upstream and on the lateral side of the cavity. Heat flux measurement has been done using indigenously developed high resistance platinum thin film gauges. High resistance gauges, as contrary to the conventionally used low resistance gauges were showing good response to the extremely low heat flux values on a flat plate with sharp leading edge. The experimental measurements of heat done on a flat plate with sharp leading edge using these gauges show good match with theoretical relation by Crabtree et al. Flow visualization using high speed camera with the cavity model and shock structures visualized were similar to reported in supersonic cavity flow. This also goes to state that in spite of the fluctuating shear layer-the main feature of hypersonic flow over a cavity ,reasonable studies can be done within the short test time of shock tunnel. Numerical Simulations by solving the Navier-Stokes equation, using the commercially available CFD package FLUENT 13.0.0 has been done to complement the experimental studies.