1.2. Definition of hypersonic flow regime:
Definition of flow regime is based on the Mach number of the flow. If Mach number is below unity then the flow is called as subsonic. Sonic flow has Mach number exactly equal to one however flow in the narrow range of Mach number beween 08-1.2 is called as transonic flow. When the flow Mach number exceeds beyond 1 then flow is called as supersonic flow. As per the thumb rules, when flow speed exceeds five times the sound speed, it is treated as hypersonic flow. However hypersonic flow has certain characteristics which when experienced in the flow, should then only be termed as hypersonic. These characteristics of hypersonic flow are mentioned below.

1. Thin Shock Layers
   Region between shock and the body (flight vehicle) is named as shock layer. From the relations between shock angle, Mach number and flow deflection angle or wedge angle, it would be clear that, for same flow deflection angle, shock angle decreases with increase in Mach number. Hence the thickness of the shock layer decreases with increase in Mach number for the same flow deflection angle. Therefore hypersonic flows have thin shock layer. This interpretation of shock layer thinness for calorically perfect gas is also applicable for thermally perfect gas and chemically reacting flow. However, complexity of flow field increases due to thin shock layer where the boundary layer thickness and shock layer thickness become comparable.
2. Entropy Layer
   

   One of the main properties of the curved shock waves in that, each streamline passing through the shock faces differential entropy rise where stronger portion of shock leads to higher entropy rise than the weaker portion. Therefore, a layer of entropy variation getting formed downstream of the shock is termed as entropy layer. Analysis of hypersonic flow becomes further troublesome with consideration of this entropy layer since according to Croco's principle the entropy layer leads to vorticity. As it was evident that the shock layer thickness decreases with increase in Mach number and shock comes closer to the sharp leading edge configurations like wedge or cone, it is also obvious that shock detachment distance decreases with increase in Mach number for blunt bodies. Hence the entropy layer exhibits strong gradient of entropy which leads to higher vorticity at higher magnitudes of Mach numbers. Due to presence of entropy layer it becomes difficult to predict the boundary layer properties and properties at the edge of the boundary layer of hypersonic flow due to interaction of boundary layer vorticity and entropy layer vorticity. This interaction is termed as vorticity interaction.

3. Viscous Interaction
   As we know, formation of boundary layer takes place near the wall due to no-slip property of the viscous fluid flow. Formation of this boundary layer takes place across enourmous loss of kinetic energy at hypersonic speeds. This kinetic energy necessarily gets converted in to thermal energy which leads to increase in temperature of the flow in the vicinity of the wall. This phenomenon is called as viscous dissipation. Viscous dissipation leads to increase in boundary layer thickness due to increase in viscosity coefficient with temperature. This situation can also be interpreted from boundary layer theory where pressure is considered to be constant across the boundary layer. This thickened boundary layer displaces outer inviscid flow hence freestream hypersonic flow encounters an inflated object which changes the shock shape and intren boundary layer parameters along with surface pressure, wall heat flux, skin friction etc. This interaction or communication loop between viscous boundary layer and outer inviscid flow is called as viscous interaction. As a result of this interaction aerodynamic parameters such as lift, drag etc deviate a lot from their base value without interaction. Hence it becomes mandatory to treat viscous interaction for hypersonic flights since whole shock layer tends to become viscous due to this interaction.
4. High-Temperature Flows
   As we know, viscous dissipation leads to higher boundary layer thickness and temperature of the boundary layer fluid. Therefore any hypersonic flight expreiences presence of high temperature fluid in the vicinity of the flight vehicle. Apart from this, blunt nosed configurations encounter very high temperatures due to normal shock present at the stagnation point. Therefore at these elevated temperaturess, treatment of fluid as calorically perfect or with constant thermodynamic properties leads to unrealistic estimations. Hence it becomes essential to take in to account the dependence of specific heats and their ratio as function of temperature for rational estimates.
   

   The dependence of thermodynamic properties on temperature mainly comes from microscopic changes in the fluid due to increase in internal energy of the fluid by the virtue of loss of kinetic energy. Increased internal energy leads initially to vibrational excitation followed by dissociation and finally ionization according to the extent of increase in internal energy. As per the order of magnitude estimate, vibrational excitation of air takes place at around 800K. Oxygen dissociation starts at around 2000 K and completes at 4000 K. At around 4000 K nitrogen dissociation commences and completes at 9000 K. Ionization of this high temperature air or mixture of gases starts from 9000 K temperature. Hence the initil air with atmosheric composition becomes plasma after 9000 K. As a result of all these reactions, hypersonic vehicle gets engulfed by reacting boundary layer and high temperature plasma. Therefore treatment of air or any fluid flowing with hypersonic speed over any configuration should be done properly by incorporating all the microscopic changes which essentially leads to change in thermodynamic properties with temperature. This dependence is highly non-linear, hence analysis or prediction of flowfield becomes tougher in this flow regime. Therefore two types of assumptions are generally made about the flow conditions for high temperature fluid as equilibrium flow and non-equilibrium flow. If the microscopic changes or reactions are at faster rate than the movement of the fluid, then it is treated as equilibrium flow otherwise it is treated as non-equilibrium flow which is difficult to analyze. All these difficulties are collectively termed as ‘real gas effects'.
   Some consequences of presence of high temperature reacting fluid or plasma in the vicinity of the flight vehicle include, influence on aerodynamic parameters, aerodynamic heating and communication block-out. Flight parameters like pitch, roll, drag, lift, defection of control surfaces get largely deviated from their usual estimate of calorically perfect gas. Presence of hot fluid near the cold vehicle surface induces heat transfer not only through convection but also through radiation. Communication waves which are necessarily radio waves get absorbed by free electrons formed from ionization of atmospheric fluid. This phenomenon is called as communication block-out where on board and ground communication gets terminated