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With a sufficiently small half-angle and properly placed center of mass, a sphere-cone can provide aerodynamic stability from Keplerian entry to surface impact.

The half-angle is the angle between the cone's axis of rotational symmetry and its outer surface, and thus half the angle made by the cone's surface edges.

The original American sphere-cone aeroshell was the Mk-2 RV reentry vehicle , which was developed in by the General Electric Corp.

The Mk-2's design was derived from blunt-body theory and used a radiatively cooled thermal protection system TPS based upon a metallic heat shield the different TPS types are later described in this article.

The Mk-2 had significant defects as a weapon delivery system, i. These defects made the Mk-2 overly susceptible to anti-ballistic missile ABM systems.

This new TPS was so effective as a reentry heat shield that significantly reduced bluntness was possible.

Subsequent advances in nuclear weapon and ablative TPS design allowed RVs to become significantly smaller with a further reduced bluntness ratio compared to the Mk Reconnaissance satellite RVs recovery vehicles also used a sphere-cone shape and were the first American example of a non-munition entry vehicle Discoverer-I , launched on 28 February The sphere-cone was later used for space exploration missions to other celestial bodies or for return from open space; e.

Space exploration sphere-cone entry vehicles have landed on the surface or entered the atmospheres of Mars , Venus , Jupiter , and Titan.

The biconic is a sphere-cone with an additional frustum attached. No accurate diagram or picture of AMaRV has ever appeared in the open literature.

However, a schematic sketch of an AMaRV-like vehicle along with trajectory plots showing hairpin turns has been published. AMaRV's attitude was controlled through a split body flap also called a split-windward flap along with two yaw flaps mounted on the vehicle's sides.

Hydraulic actuation was used for controlling the flaps. AMaRV was guided by a fully autonomous navigation system designed for evading anti-ballistic missile ABM interception.

Non- axisymmetric shapes have been used for manned entry vehicles. One example is the winged orbit vehicle that uses a delta wing for maneuvering during descent much like a conventional glider.

Objects entering an atmosphere from space at high velocities relative to the atmosphere will cause very high levels of heating.

Reentry heating comes principally from two sources: [14]. As velocity increases, both convective and radiative heating increase.

At very high speeds, radiative heating will come to quickly dominate the convective heat fluxes, as convective heating is proportional to the velocity cubed, while radiative heating is proportional to the velocity exponentiated to the eighth power.

Radiative heating—which is highly wavelength dependent—thus predominates very early in atmospheric entry while convection predominates in the later phases.

At typical reentry temperatures, the air in the shock layer is both ionized and dissociated. There are four basic physical models of a gas that are important to aeronautical engineers who design heat shields:.

Almost all aeronautical engineers are taught the perfect ideal gas model during their undergraduate education. Most of the important perfect gas equations along with their corresponding tables and graphs are shown in NACA Report The perfect gas theory is elegant and extremely useful for designing aircraft but assumes that the gas is chemically inert.

From the standpoint of aircraft design, air can be assumed to be inert for temperatures less than K at one atmosphere pressure.

The perfect gas theory begins to break down at K and is not usable at temperatures greater than 2, K. For temperatures greater than 2, K, a heat shield designer must use a real gas model.

An entry vehicle's pitching moment can be significantly influenced by real-gas effects. Both the Apollo CM and the Space Shuttle were designed using incorrect pitching moments determined through inaccurate real-gas modelling.

The Apollo-CM's trim-angle angle of attack was higher than originally estimated, resulting in a narrower lunar return entry corridor.

The actual aerodynamic center of the Columbia was upstream from the calculated value due to real-gas effects. Young and Robert Crippen had some anxious moments during reentry when there was concern about losing control of the vehicle.

An equilibrium real-gas model assumes that a gas is chemically reactive, but also assumes all chemical reactions have had time to complete and all components of the gas have the same temperature this is called thermodynamic equilibrium.

When air is processed by a shock wave, it is superheated by compression and chemically dissociates through many different reactions.

Direct friction upon the reentry object is not the main cause of shock-layer heating. It is caused mainly from isentropic heating of the air molecules within the compression wave.

Friction based entropy increases of the molecules within the wave also account for some heating. An approximate rule of thumb for shock wave standoff distance is 0.

One can estimate the time of travel for a gas molecule from the shock wave to the stagnation point by assuming a free stream velocity of 7.

This is roughly the time required for shock-wave-initiated chemical dissociation to approach chemical equilibrium in a shock layer for a 7. Consequently, as air approaches the entry vehicle's stagnation point, the air effectively reaches chemical equilibrium thus enabling an equilibrium model to be usable.

For this case, most of the shock layer between the shock wave and leading edge of an entry vehicle is chemically reacting and not in a state of equilibrium.

The Fay-Riddell equation , [10] which is of extreme importance towards modeling heat flux, owes its validity to the stagnation point being in chemical equilibrium.

The time required for the shock layer gas to reach equilibrium is strongly dependent upon the shock layer's pressure. Determining the thermodynamic state of the stagnation point is more difficult under an equilibrium gas model than a perfect gas model.

Under a perfect gas model, the ratio of specific heats also called isentropic exponent , adiabatic index , gamma , or kappa is assumed to be constant along with the gas constant.

For a real gas, the ratio of specific heats can wildly oscillate as a function of temperature. Under a perfect gas model there is an elegant set of equations for determining thermodynamic state along a constant entropy stream line called the isentropic chain.

For a real gas, the isentropic chain is unusable and a Mollier diagram would be used instead for manual calculation.

However, graphical solution with a Mollier diagram is now considered obsolete with modern heat shield designers using computer programs based upon a digital lookup table another form of Mollier diagram or a chemistry based thermodynamics program.

The chemical composition of a gas in equilibrium with fixed pressure and temperature can be determined through the Gibbs free energy method.

Gibbs free energy is simply the total enthalpy of the gas minus its total entropy times temperature. A chemical equilibrium program normally does not require chemical formulas or reaction-rate equations.

The program works by preserving the original elemental abundances specified for the gas and varying the different molecular combinations of the elements through numerical iteration until the lowest possible Gibbs free energy is calculated a Newton-Raphson method is the usual numerical scheme.

The data base for a Gibbs free energy program comes from spectroscopic data used in defining partition functions.

CEA is quite accurate up to 10, K for planetary atmospheric gases, but unusable beyond 20, K double ionization is not modelled.

CEA can be downloaded from the Internet along with full documentation and will compile on Linux under the G77 Fortran compiler. A non-equilibrium real gas model is the most accurate model of a shock layer's gas physics, but is more difficult to solve than an equilibrium model.

As of [update] , the simplest non-equilibrium model was the Lighthill-Freeman model. N 2 dissociation and recombination. Because of its simplicity, the Lighthill-Freeman model is a useful pedagogical tool, but is unfortunately too simple for modelling non-equilibrium air.

Air is typically assumed to have a mole fraction composition of 0. The five species model assumes no ionization and ignores trace species like carbon dioxide.

When running a Gibbs free energy equilibrium program, [ clarification needed ] the iterative process from the originally specified molecular composition to the final calculated equilibrium composition is essentially random and not time accurate.

With a non-equilibrium program, the computation process is time accurate and follows a solution path dictated by chemical and reaction rate formulas.

The five species model has 17 chemical formulas 34 when counting reverse formulas. The Lighthill-Freeman model is based upon a single ordinary differential equation and one algebraic equation.

The five species model is based upon 5 ordinary differential equations and 17 algebraic equations. The five species model is only usable for entry from low Earth orbit where entry velocity is approximately 7.

The five-species model is no longer accurate and a twelve-species model must be used instead. An important aspect of modelling non-equilibrium real gas effects is radiative heat flux.

If a vehicle is entering an atmosphere at very high speed hyperbolic trajectory, lunar return and has a large nose radius then radiative heat flux can dominate TPS heating.

Radiative heat flux during entry into an air or carbon dioxide atmosphere typically comes from asymmetric diatomic molecules; e.

These molecules are formed by the shock wave dissociating ambient atmospheric gas followed by recombination within the shock layer into new molecular species.

The newly formed diatomic molecules initially have a very high vibrational temperature that efficiently transforms the vibrational energy into radiant energy; i.

The whole process takes place in less than a millisecond which makes modelling a challenge. The experimental measurement of radiative heat flux typically done with shock tubes along with theoretical calculation through the unsteady Schrödinger equation are among the more esoteric aspects of aerospace engineering.

Most of the aerospace research work related to understanding radiative heat flux was done in the s, but largely discontinued after conclusion of the Apollo Program.

Radiative heat flux in air was just sufficiently understood to ensure Apollo's success. However, radiative heat flux in carbon dioxide Mars entry is still barely understood and will require major research.

The frozen gas model describes a special case of a gas that is not in equilibrium. The name "frozen gas" can be misleading. A frozen gas is not "frozen" like ice is frozen water.

Rather a frozen gas is "frozen" in time all chemical reactions are assumed to have stopped. Chemical reactions are normally driven by collisions between molecules.

If gas pressure is slowly reduced such that chemical reactions can continue then the gas can remain in equilibrium.

However, it is possible for gas pressure to be so suddenly reduced that almost all chemical reactions stop.

For that situation the gas is considered frozen. The distinction between equilibrium and frozen is important because it is possible for a gas such as air to have significantly different properties speed-of-sound, viscosity etc.

Frozen gas can be a significant issue in the wake behind an entry vehicle. During reentry, free stream air is compressed to high temperature and pressure by the entry vehicle's shock wave.

Non-equilibrium air in the shock layer is then transported past the entry vehicle's leading side into a region of rapidly expanding flow that causes freezing.

The frozen air can then be entrained into a trailing vortex behind the entry vehicle. Correctly modelling the flow in the wake of an entry vehicle is very difficult.

Thermal protection shield TPS heating in the vehicle's afterbody is usually not very high, but the geometry and unsteadiness of the vehicle's wake can significantly influence aerodynamics pitching moment and particularly dynamic stability.

A thermal protection system , or TPS, is the barrier that protects a spacecraft during the searing heat of atmospheric reentry.

A secondary goal may be to protect the spacecraft from the heat and cold of space while in orbit. Multiple approaches for the thermal protection of spacecraft are in use, among them ablative heat shields, passive cooling, and active cooling of spacecraft surfaces.

The ablative heat shield functions by lifting the hot shock layer gas away from the heat shield's outer wall creating a cooler boundary layer.

The boundary layer comes from blowing of gaseous reaction products from the heat shield material and provides protection against all forms of heat flux.

The overall process of reducing the heat flux experienced by the heat shield's outer wall by way of a boundary layer is called blockage.

Ablation occurs at two levels in an ablative TPS: the outer surface of the TPS material chars, melts, and sublimes , while the bulk of the TPS material undergoes pyrolysis and expels product gases.

The gas produced by pyrolysis is what drives blowing and causes blockage of convective and catalytic heat flux. Pyrolysis can be measured in real time using thermogravimetric analysis , so that the ablative performance can be evaluated.

Radiative heat flux blockage was the primary thermal protection mechanism of the Galileo Probe TPS material carbon phenolic. Carbon phenolic was originally developed as a rocket nozzle throat material used in the Space Shuttle Solid Rocket Booster and for reentry-vehicle nose tips.

Ames Research Center was ideal, since it had numerous wind tunnels capable of generating varying wind velocities.

Initial experiments typically mounted a mock-up of the ablative material to be analyzed within a hypersonic wind tunnel. Many spacecraft thermal protection systems have been tested in this facility, including the Apollo, space shuttle, and Orion heat shield materials.

The thermal conductivity of a particular TPS material is usually proportional to the material's density.

If the heat flux experienced by an entry vehicle is insufficient to cause pyrolysis then the TPS material's conductivity could allow heat flux conduction into the TPS bondline material thus leading to TPS failure.

Consequently, for entry trajectories causing lower heat flux, carbon phenolic is sometimes inappropriate and lower-density TPS materials such as the following examples can be better design choices:.

For Viking-1, the TPS acted as a charred thermal insulator and never experienced significant ablation. Viking-1 was the first Mars lander and based upon a very conservative design.

The Viking aeroshell had a base diameter of 3. SLAV is applied by packing the ablative material into a honeycomb core that is pre-bonded to the aeroshell's structure thus enabling construction of a large heat shield.

Phenolic-impregnated carbon ablator PICA , a carbon fiber preform impregnated in phenolic resin , [26] is a modern TPS material and has the advantages of low density much lighter than carbon phenolic coupled with efficient ablative ability at high heat flux.

It is a good choice for ablative applications such as high-peak-heating conditions found on sample-return missions or lunar-return missions.

PICA's thermal conductivity is lower than other high-heat-flux-ablative materials, such as conventional carbon phenolics. Stardust's heat shield 0.

It was first flight tested on the Crew Dragon spacecraft in during the flight demonstration mission , in April , and put into regular service on that spacecraft in The BIP was at the attachment points between the aeroshell's backshell also called the afterbody or aft cover and the cruise ring also called the cruise stage.

SIRCA is a monolithic, insulating material that can provide thermal protection through ablation. It is the only TPS material that can be machined to custom shapes and then applied directly to the spacecraft.

There is no post-processing, heat treating, or additional coatings required unlike Space Shuttle tiles. Since SIRCA can be machined to precise shapes, it can be applied as tiles, leading edge sections, full nose caps, or in any number of custom shapes or sizes.

NASA originally used it for the Apollo capsule in the s, and then utilized the material for its next-generation beyond low-Earth-orbit Orion spacecraft, slated to fly in the late s.

Thermal soak is a part of almost all TPS schemes. For example, an ablative heat shield loses most of its thermal protection effectiveness when the outer wall temperature drops below the minimum necessary for pyrolysis.

From that time to the end of the heat pulse, heat from the shock layer convects into the heat shield's outer wall and would eventually conduct to the payload.

An LI tile exposed to a temperature of 1, K on one side will remain merely warm to the touch on the other side. However, they are relatively brittle and break easily, and cannot survive in-flight rain.

In some early ballistic missile RVs e. However, the earlier version of this technique required a considerable quantity of metal TPS e.

Modern designers prefer to avoid this added mass by using ablative and thermal-soak TPS instead.

Thermal protection systems relying on emissivity use high emissivity coatings HECs to facilitate radiative cooling , while an underlying porous ceramic layer serves to protect the structure from high surface temperatures.

High thermally stable emissivity values coupled with low thermal conductivity are key to the functionality of such systems. Radiatively cooled TPS can be found on modern entry vehicles, but reinforced carbon—carbon RCC also called carbon—carbon is normally used instead of metal.

This high temperature made carbon an obvious choice as a radiatively cooled TPS material. Disadvantages of RCC are that it is currently expensive to manufacture, is heavy, and lacks robust impact resistance.

Some high-velocity aircraft , such as the SR Blackbird and Concorde , deal with heating similar to that experienced by spacecraft, but at much lower intensity, and for hours at a time.

Studies of the SR's titanium skin revealed that the metal structure was restored to its original strength through annealing due to aerodynamic heating.

These TPS materials are based on zirconium diboride and hafnium diboride. SHARP TPS materials enable sharp leading edges and nose cones to greatly reduce drag for airbreathing combined-cycle-propelled spaceplanes and lifting bodies.

They are structurally stronger than RCC, and, thus, do not require structural reinforcement with materials such as Inconel. Various advanced reusable spacecraft and hypersonic aircraft designs have been proposed to employ heat shields made from temperature-resistant metal alloys that incorporate a refrigerant or cryogenic fuel circulating through them, and one such spacecraft design is currently under development.

Such a TPS concept was proposed [ when? SpaceX is currently developing an actively cooled heat shield for its Starship spacecraft where a part of the thermal protection system will be a transpirationally cooled outer-skin design for the reentering spaceship.

In the early s various TPS systems were proposed to use water or other cooling liquid sprayed into the shock layer, or passed through channels in the heat shield.

Advantages included the possibility of more all-metal designs which would be cheaper to develop, be more rugged, and eliminate the need for classified technology.

The disadvantages are increased weight and complexity, and lower reliability. The concept has never been flown, but a similar technology the plug nozzle [47] did undergo extensive ground testing.

In , aircraft designer Burt Rutan demonstrated the feasibility of a shape-changing airfoil for reentry with the sub-orbital SpaceShipOne.

The wings on this craft rotate upward into the feather configuration that provides a shuttlecock effect. Thus SpaceShipOne achieves much more aerodynamic drag on reentry while not experiencing significant thermal loads.

The configuration increases drag, as the craft is now less streamlined and results in more atmospheric gas particles hitting the spacecraft at higher altitudes than otherwise.

The aircraft thus slows down more in higher atmospheric layers which is the key to efficient reentry. Secondly, the aircraft will automatically orient itself in this state to a high drag attitude.

However, the velocity attained by SpaceShipOne prior to reentry is much lower than that of an orbital spacecraft, and engineers, including Rutan, recognize that a feathered reentry technique is not suitable for return from orbit.

On 4 May , the first test on the SpaceShipTwo of the feathering mechanism was made during a glideflight after release from the White Knight Two.

Premature deployment of the feathering system was responsible for the VSS Enterprise crash , in which the aircraft disintegrated, killing the co-pilot.

It may be desirable to combine lifting and nonlifting entry in order to achieve some advantages For landing maneuverability it obviously is advantageous to employ a lifting vehicle.

The total heat absorbed by a lifting vehicle, however, is much higher than for a nonlifting vehicle Nonlifting vehicles can more easily be constructed The larger the device, the smaller is the heating rate.

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