Mach 7 Mph - In aerodynamics, hypersonic speed is that which exceeds 5 times the speed of sound, often starting at speeds of Mach 5 and above.
The exact Mach number that can be said to fly at hypersonic speed varies, because individual physical changes in the air flow (such as molecular fragmentation and ionization) occur at different speeds; These effects collectively become significant around Mach 5-10. A hypersonic system may be defined as a motion in which the specific heat capacity varies with the temperature of the flow as the kinetic energy of the moving object is converted into heat.
Mach 7 Mph
Although the definition of hypersonic flow can be somewhat vague and generally debatable (mainly because of the discontinuity between supersonic and hypersonic flow), hypersonic flow can be characterized by a number of physical forms that are now very similar to supersonic flow. .
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As the Mach number of the body increases, the density behind the bow shock produced by the body also increases, which corresponds to a decrease in the volume behind the shock due to conservation of mass. Consequently, the distance between the bow shock and the body decreases at higher Mach numbers.
As the Mach number increases, the tropopause across the shock also increases, resulting in a strong tropopause gradient and high vortical flow present in the boundary layer.
At high Mach numbers, part of the large kinetic energy associated with the flow is converted into internal energy in the fluid due to viscous effects. An increase in internal energy is felt as an increase in temperature. Since the pressure gradient for the flow in the boundary layer is almost zero for normal to low to moderate hypersonic Mach numbers, the increase in temperature through the boundary layer corresponds to a decrease in density. This causes the bottom of the boundary layer to expand, so that the boundary layer over the body thickens and can join the shock wave near the leading edge of the body.
Due to the manifestation of viscous dissipation, high-temperature non-equilibrium chemical flow properties such as vibrational excitation and molecular dissociation and ionization lead to convective and radiative heat fluxes.
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This section should be expanded with: Added "Jeral spaceship characteristics" to aircraft characteristics in the table. You can help by donating. (June 2021)
Although "subsonic" and "supersonic" usually refer to speeds below and above the local speed of sound, however, aerodynamicists often use these terms to refer to a certain range of Mach values. This occurs because there is a "transonic regime" around M=1 where the approximate Navier-Stokes equation used for subsonic design no longer applies, in part because the flow is localized locally at M=1. beyond when freestream.
The "supersonic regime" generally refers to the range of Mach numbers for which linear theory can be used; For example, where the flow (air) does not react chemically and where the heat transfer between the air and the vehicle can be reasonably neglected in the calculations. In general, NASA defines "high" hypersonics as Mach numbers from 10 to 25, and redefined speeds as anything above Mach 25. The Soyuz and Dragon space capsules are returning from spacecraft operating in the system; the first space shuttle to operate; various reusable spacecraft under development such as SpaceX Starship and Rocket Lab Electron; as well as spaceships (theoretically).
In the table below, reference is made to "regulation" or "Mach value range" instead of the usual meanings of "subsonic" and "supersonic".
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Most commercial and propeller-driven turbofan aircraft have high aspect ratio (slim) wings, and rounded features such as noses and edges.
Transonic aircraft almost always have swept wings to delay drag-diverge, supercritical wings to delay the onset of wave drag, and often feature designs that follow the principles of Whitcombe's field rule.
Aircraft designed to fly at supersonic speeds exhibit large differences in aerodynamic design due to radical differences in fluid flow behavior above Mach 1. Sharp edges, thin airfoil sections, and moving tailplanes/canards are all common. Modern fighter jets have to make compromises to maintain low speed.
Cold nickel or titanium skin; Due to the dominance of the interface effect, the design is highly integrated, instead of being assembled from components designed separately, where a small change in one component can cause a large change in the airflow in all the other components, which can affect the behavior. The result is that no single component can be designed without knowing how all the other components will affect the air flow around the ship, and changes in one component will affect all the other components simultaneously. A redesign may be necessary.
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; small wings See Boeing X-51 Waverider, BrahMos-II, X-41 General Aero Vehicle, DF-ZF, Hypersonic Technology Demonstrator Vehicle, Hypersonic Airborne Weapon Concept (HAWC, pronounced Hawk), Shaurya Missile.
Thermal control is a major design consideration. The structure must be designed to run hot, or protected by a special silicate box or similar. Chemically reactive flows can also cause corrosion on vehicle skins, with free atomic oxygen characteristic of very high velocity flows. Examples include the 53T6 (Mach 17), Hypersonic Technology Vehicle 2 (Mach 20), LGM-30 Minuteman (Mach 23), Agni-V (Mach 24), DF-41 (Mach 25), and Avangard (Mach 20-27) . ) including ). Hypersonic designs are often forced into blunter configurations due to increased aerodynamic heating with reduced radius of curvature.
Airflow classification relies on some similarity criteria, which allows for the simplification of an almost infinite number of test cases into similar groups. For transonic and incompressible flows, Mach and Reynolds numbers alone allow a good classification of many flow cases.
However, hypersonic flow requires other uniformity parameters. First, the analytical equation for the oblique shock angle becomes incomplete in the Mach number approximation at high (~>10) Mach numbers. Second, the strong formation of shocks around an aerodynamic body means that the freestream Reynolds number is less useful as a predictor of the behavior of the boundary layer on the body (although it is still important). Finally, the increased temperature of hypersonic flow means that real gas effects become important. Research in hypersonics is often referred to as aerothermodynamics rather than aerodynamics.
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The introduction of real gas effects means that more variables are needed to describe the complete state of the gas. While a static gas can be described by three variables (pressure, temperature, adiabatic index), and a moving gas by four (flow rate), even hot gases in chemical equilibrium require an equation of state for the chemical components of the gas. , and the non-equilibrium gas solves the equation of state using time as an additional variable. This means that for non-equilibrium flow, something between 10 and 100 variables may be needed to describe the state of the gas at any given time. Furthermore, rare hypersonic flows (usually defined by Knuds numbers above 0.1) do not obey the Navier–Stokes equations.
Hypersonic flow is usually classified by total energy, total enthalpy (MJ/kg), total pressure (kPa-MPa), stagnation pressure (kPa-MPa), stagnation temperature (K), or flow velocity (km/s).
Wallace D. Hayes developed the similarity parameter of Whitcomb's Area Rule, which allows comparison of similar structures.
Hypersonic flow can be divided into several regimes. The choice of the regime is harsh, because of the blurred boundaries where certain effects can be found.
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In this regime, the gas can be considered as an ideal gas. In this regime the flow is still Mach number depdt. The simulation begins with immersion using a constant temperature wall, instead of the adiabatic wall usually used at low speed. The lower limit of this region is around Mach 5, where ramjets become ineffective, and the upper limit is around Mach 10–12.
This is a subset of the perfect gas system, where the gas can be assumed to be chemically perfect, but the rotational and vibrational temperatures of the gas must be considered separately, leading to a two-temperature model. See especially the modeling of supersonic nozzles, where vibrational freezing becomes important.
In this system, diatomic or polyatomic gases (mostly gases found in the atmosphere) begin to dissociate when they are exposed to arc shocks produced by the body. The surface catalyst plays a role in the surface heating calculation, which means that the type of surface material also has an effect on the flux. The lower limit of this system is the part of the gas mixture that first begins to separate at the flow stagnation point (about 2000 K for nitrogen). At the upper limit of this regime, ionization effects begin to affect the flow.
In this regime the population of ionized electrons from the steady flow becomes important, and the electrons must be modeled separately. Often the electron temperature is considered separately from the temperature of other gas components. This region is for freestream flow velocities of around 3–4 km/s. The gas in this region is modeled as a non-radiating plasma.
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Above about 12 km/s, the heat transfer in the vehicle changes from conductive to radiative dominated. Gas modeling in this system is divided into two categories:
Modeling optically dense gases is very difficult, because, by calculating the radiation at each point,
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