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reentry of space vehicle full report
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Re-entry capsules promises to intensify international competition in launch services, microgravity research and space technology development. These systems will also confer an important strategic advantage in the conduct of materials and in life science research.
The objective of this paper is to provide a modest degree of understanding of the complex inter-relation which exist between performance requirements mission constraints , vehicle design and trajectory selection of typical re-entry mission. A brief presentation of the flight regimes, the structural loading and heating environment experienced by booth no lifting and lifting re-entry vehicle is given.

The successful exploration of space requires a system that will reliably transport payload such as personnel and instrumental etc. into space and return them back to earth without subjecting them an uncomfortable or hazardous environment. In other words, the spacecraft and its payloads have to be recovered safely into the earth. We have seen the re-entry capsules and winged space vehicles approach the earth followed by safe landing. However, this could be accomplished only after considerable research in high speed aerodynamics and after many parametric studies to select the optimum design concept.

Re-entry systems were among the first technologies developed in 1960s for military photo-reconnaissance, life science and manned space flights. By 1970s, it led to the development of new refurbish able space shuttles. Today space technology has developed to space planes which intend to go and come back regularly from earth to space stations. USAâ„¢s HERMS and Japanâ„¢s HOPE is designed to land at conventional airports. Few significant advances in current proposed re-entry capsules are ballistic designs to reduce development and refurbishable cost, to simplify operations.
For entering into atmospheric and non-atmospheric planet the problem involves is reducing the spacecraftâ„¢s speed . For an atmospheric planet the problem involves essentially deceleration, aerodynamic heating, control of time & location of landing. For non-atmospheric planets, the problem involves only deceleration and control of time & location of landing.
The vehicle selected to accomplish a re-entry mission incorporates a thick wing , subsonic ( Mach < 1 ) airfoil modified to meet hypersonic (Mach>> 1 ) thermodynamic requirements. The flight mechanics of this vehicle are unique in that rolling manoeuvres are employed during descent such that dynamic loading and aerodynamic heating are held to a minimum.
Therefore re-entry technology requires studies in the following areas:
1. Deceleration
2. Aerodynamic heating & air loads
3. Vehicle stability
4. Thermal Protection Systems (TPS)
5. Guidance and Landing.

The safe recovery of the spacecraft and its payloads is made possible by the re-entry mission. According to the different constraints the mission profile can be divided into three distinct flight segments:-
1. Deorbit and Descent to sensible atmosphere at an altitude of nearly 120kms.
2. Re-entry and hypersonic glide fight.
3. Transition flight phase, final approach and landing.
The unguided first flight segment (Keplarian trajectory) initiated by a rocket deboost maneuver at a specific orbital point determines the flight condition at re-entry. The second flight segment covers the atmospheric glide at an altitude of 120 km to 30 km during which the re-entry vehicleâ„¢s high initial kinetic energy is dissipated by atmospheric breaking. The third flight segment does the final approach and landing. All these phases are shown in Fig.1.
The various forces acting on the re-entry vehicle are:-
1. Gravitational force acting towards the centre of the planet.
2. Gas dynamic force opposite to the direction of motion of the vehicle.
3. Centrifugal and gas dynamic lift force acting normal to the direction of
4. motion of the vehicle.
Along the re-entry flight several mission constraints much be imposed arising from the structural limit, crew comfort and control limits.
These limits require the flight state of the vehicle to the constrained such that the:-
1. Load factor n <= n max
2. Dynamic pressure q = 1/2 pv2 <= qmax
3. Heat flux Q <= Qmax
4. Heat load Q1 = O.T Qdt <= Q1max
5. Surface temperature T <= Tmax
The maximum admissible values of these factors are highly dependent on the state of the technology involved regarding heat resistant , light weight materials and structures.
The actual flight loads experienced by the vehicle depends upon:
1. Local atmospheric environment (eg: density, temperature)
2. Current flight static conditions (eg : velocity, angle of attack)
3. Vehicle properties(e.g.: geometry, weight, aerodynamics) and thus specific re-entry trajectories and design parameters.
The most important design parameters with respect to re-entry performance are given by :
1. Wing loading, m /A (kg /m2)
2. Ballistic coefficient, B= m /(Cd * A)
3. Lift to Drag ratio, L / D
, where L= aerodynamic lift
A= aerodynamic lift area
M= vehicle mass
D = drag
Cd= drag coefficient
Depending on the specified mission requirements the second or third property is chosen as design drivers.

Parts of a space vehicle

An entry corridor is a range of entry conditions within which an entry is possible. The Ëœundershootâ„¢ boundary and Ëœovershootâ„¢ boundary forms the upper and the lower limits of the entry corridor. Terrestrial flights are tolerant of guidance error accompanying a landing approach. An undershoot may cause destruction of vehicle during entry and an undershoot may result in a homeless exit to space.
Figure shows the explanation of entry corridor and possible path for vehicle with lift to Venus, Mars, and Titan.
If the guidance error results in an excessive undershoot as shown by the two dashed trajectories, the vehicle will enter the atmosphere at an excessively steep angle, thereby experiencing too much deceleration. If the guidance error results in an excessive overshoot as shown by the two outer dashed trajectories, the vehicle will not slow down considerably in order to complete entry in a single pass. Hence the shaded portions representing excessively overshoot and undershoot are excluded as not representing the intended entry manoeuvre.

Although overshoot at hyperbolic velocity > (2gRo)1/2 may result in a homeless exit to space, overshoot at the outer corridor at parabolic speed = (2gRo)1/2 or at an elliptical speed < (2gRo)1/2 or at an elliptical speed , (2gRo)1/2 will result in a multipass entry.

Overshoot passage was considered a good way to came back from moon or more distant planets. Each pass through the atmosphere would slow down the vehicle a little , so that it would return in a series of successively shorter ellipses. In this way the heat problem would be solved, heat taken on each approach being radiated during the next outward journey.
When either truly circular entry is made inside entry corridor or final entry is made through multipass scheme, the descent trajectory through the atmosphere is similar to entry from a satellite orbit, namely ballistic path, glide path or slip path.
The width of entry corridor (rp) for non-lifting (L/D<0) and lifting (L/D>1) entry into various planetary atmosphere is given in the table below for entry at planetary velocity (2gro)1/2.

Corridor width
5g limit 10g limit
L/D=0 L/D=1 L/D=1
modulated L/D=0 L/D=1 L/D=1
Venus 0 43 58 13 84 113
Earth 0 43 55 11 82 105
Mars 338 482 595 644 885 1159
Jupiter 0 55 68 0 84 113

The most important problem that a re-entry mission has to face is the atmospheric deceleration. The deceleration forces can be as great as 600 to 900N for unmanned space probes. Space shuttles use their wings to skim the atmosphere and stretch the slow down period to more than 15 minutes and thereby reducing the deceleration forces to about 15N.

When too much deceleration is intolerable, aerodynamic lift must be used. Lift can reduce rate of descent, thus lengthening the path to the ground and decreasing the maximum deceleration. Before entering to the atmosphere, the motion of a vehicle is governed by its own inertia and its gravitational force. When the vehicle enters the atmosphere the gas dynamic force modify such motion. Gas dynamic force acts in a direction opposite to the vehicle motion.
Figure gives an idea about the drag and lift forces. The gas dynamic drag force causes the reduction in the vehicles velocity and the centrifugal and lift forces cause acceleration normal to the direction of motion. The gas dynamic lift and drag motion and the resultant acceleration and the deceleration very directly with the atmospheric density P and square of velocity V2 . Thus vehicleâ„¢s deceleration varies with PV2.
As the vehicle approaches the planet, it first encounters an atmosphere of very low density, as it goes deeper and deeper density of the atmosphere increases rapidly and velocity begin to decrease due to drag. The deceleration force is the product of two quantities, one increasing and the other decreasing. Initially deceleration increases and at some point of velocity begins to decrease more rapidly than the density increase, resulting in a minimum deceleration with subsequent decreasing acceleration. Figure shows the changes during re-entry.
When the spacecraft has lost most of its speed, it falls freely through air. Parachutes slow it further down and a small rocket is fired in the final seconds of descent to soften the impact of landing. For approach to non-atmospheric planets, the absence of aerodynamic lift and drag necessitates the use of reverse thrust rockets to slow down the vehicle for safe landing. Because of the absence of atmosphere, there is no problem of aerodynamic heating. Programming of rocket thrust is desired for best controlling the time and location of landing.
The major concern of re-entry is to find a way to survive the aerodynamic heating. This obstacle was named as the ËœThermal Barrierâ„¢. A vehicle approaching the earth or a planetary atmosphere from space or from orbit possesses a large amount of kinetic energy due to its speed and potential energy by virtue of its position. These energies have to be dissipated and converted into heat to decelerate the vehicle to zero velocity and altitude. From speed of 8 Km/s to energy per mass of 31.4 MJ/Kg must be dissipated.
At high speed associated with re-entering from space, air cannot flow out of the way of the on rushing spacecraft fast enough. When the vehicle encounters the atmosphere a shock will from ahead of the nose of the vehicle heating the atmosphere in this region to a very high temperature. As the vehicle plunges into deeper and denser atmosphere the vehicle will increasingly be heated by the enveloping layer of incandescent atmosphere, while the speed of the vehicle will continuously be reduced by the braking force of the atmosphere. In this manner the vehicleâ„¢s K.E is converted into heat. If all the vehicleâ„¢s energy were converted to heat within the vehicle itself, it would be more than enough to vaporize the vehicle.
There are two ways in which the total energy is dissipated from the vehicle. They are:
i. By waves unloading major part of the heat on the atmosphere by the shock waves.
ii. To radiate heat away from hot surface of the vehicle.
Figure shows the shock wave formed fir blunt and streamlined configurations.
The diversion of heat by strong shock waves is the result of molecular interaction in the gas around the vehicle. When molecules strike the forward surface they bounce back. Many of the rebounding particles collide with the oncoming molecules diverting them from the surface and preventing them by heating it by direct impact. A blunt nose produces the strongest shock wave. Since the fraction of the total heat load that is transferred is directly proportional to the strength of the shock waves the stronger the shock wave, the smaller the frictional component introduced into the body. For the reason, the entry vehicle is given a blunt shape rather than a streamlined configuration that has been elastic solution to aerodynamic heating at lower speed.
Thus, one or more of the following methods can disperse the heat reaching the vehicle:
a. Making skin material thick enough to act as a sink.
b. Radiation cooled shield with a thin metal skin and negligible heat sink capacity may be used. Here the sink temperature in radiation equilibrium situation must not exceed the heat capacity of the metal.
c. To use a heat shield constructed of inoculated layer of fiberglass and similar materials. Under intense heat the outer layer of the shield chars melts and vaporizes (Ablation).
Another mode of radiating heat is by undergoing multiple passes through the atmosphere; heat developed on each approach is radiated away in the next loop. Insulating plates of quartz fibre glued to skin creates a heat shield that protects against fierce heat.

Strength of shock wave determines the heat and drag applied directly to the vehicle by air friction. Slender shape creates weak shock wave and therefore a heavy frictional layer. Blunt has strong shock and little friction.

The prime requirements during the selection of materials is for minimizing the mass which can achieve its goals during the dynamic loading with which it is presented during the testing and launching phases and finally in zero gravity operational environment. The selection of appropriate material for an application requires knowledge of the way each property can best be used and where each limitation must be recognized.
Selection can encompass the following:
¢ Specific strength
¢ Stiffness
¢ Stress concentration resistance
¢ Fracture and fatigue resistance
¢ Thermal parameters
¢ Ease of manufacture modification.
The materials usually used for the space vehicle structures are Aluminium alloy, Magnesium alloy, Titanium alloy, Beryllium alloy, Ferrous alloy, Stainless Steel and composite materials. These alloys are probably made on the basis of maximum allowable temperature limit. Hollow and reduced section structural members such as tube and beams can exhibit stiffness characteristics compared with solid bars. Honey comb sections may be used to create panels with extremely low weight with very high stiffness. The property of honey comb panel is that the weight per unit moment of inertia is proportional to the thickness of the face times the density of material. Nickel alloy, Titanium and Aluminium are the three probable candidates for a re-entry vehicle structure.
Table-2 shows the commonly used materials for space probes with its mechanical properties.

Thermal protection system is of considerable importance in system studies. The principal type of systems used depends on either or both of the essential properties.
¢ Absorption of the heat by the surface materials by the temperature rise, phase change or chemical change.
¢ Rejection of heat by mass efflux from the surface and or radiation.
These take the following five forms.
i. Radiate, cooling from the surface of insulating materials supported on relatively cool and convectional structures.
ii. Transpiration cooling
iii. Use of vehicle structure as heat sink
iv. Ablative cooling materials supported on insulation
v. Radiative cooling for high temperature structural materials.
The simplest and lightest weight is scheme-C. The techniques by which the primary structure is protected from thermal environment is a radiation cooled outer surface which is backed by a light weight high temperature insulating material.
The main advantages of this scheme are:
1. Under ordinary conditions the entry system should have multimission capability. The outer shield will not ablate under normal conditions and thus keep the structure relatively cooled.
2. In the event of failure, such an error in the entry angle or entry from an aborted flight occurs and results in a slower thermal environment, the insulation exhibits ablating characteristics and absorbs sufficient amount of heat to protect the structure.
Scheme-A shows the combinations thick skin and radiation cooling in which radiations as well as sandwiches by low density thermal conductive material cooling structure.
Scheme-B shows transpiration cooling system which consists of a porous structure. It is equipped with a fluid supply where the fluid reaches the top of the structural surface by capillary action and thus keeps the surface cool.
Scheme-C illustrates the ablative insulating system in which the ablative coating melts and vaporizes due to high temperature thus saving the structure from melting.
Scheme-D is known as the non-ablative system. The outer surface of this insulating system is a material with good high temperature properties. The outer surface is under with one of the two low density conductivity product [PK] internal insulating materials. The outer of the two insulations has somewhat higher values of [PK] but has better higher temperature resistance. The thermal gradient through the outer layer of internal insulation allows the lower insulation to operate with its temperature limits. The composite system has been tested with interface up to 1000F with good results.

Among the several heat protection methods the heat sink method employing a high heat capacity metals such as Beryllium Oxide can absorb about 6.3 MJ/Kg without significant erosion. For high heating rates the ablation principle provides an efficient heat protection method. In this case, the metal is allowed to melt and vaporize and hence this thermal protective system is not reusable. Ablative heat shields are often impregnated with carbon fibres, which provides structural reinforcement and form a char layer that radiates heat. Reasonable TPS candidates such as ceramic tiles used for shuttle orbiter or the ceramic shingle/stand off concept developed for HERMES employ radiation cooling. The vertical plane and lateral direction can be modulated to limit the flight loads to compensate for non-nominal entry for guiding the vehicle to a restricted recovery area. Fig.7 shows the variations of heating rates and exposure times for various generic vehicles.

C.S view of the wing of the space vehicle

Landing is the final stage of re-entry mission. The accuracy requirements of the guidance and control system are largely required for re-entry phase of mission. The guidance system requirements for re-entry are a function of the following.
1. Initial condition errors at the re-entry interface.
2. Initial stable number alignment accuracy.
3. Inertial component accuracy.
4. Guidance equation mechanization.
5. Re-entry requirements.
6. Allowable re-entry dispersion.
Landing phases are different for capsule and winged glide vehicles. In the case of the capsules, most of its speed, it falls freely through air. Parachutes slow it further and small rockets are fired in the final seconds of descent to soften the impact of landing. The oceans are the prime landing areas. Water landing do not require the high precision of landing area. The early US space capsules used the cushioning of water and splashed down into oceans.
The major advantage in the case of winged glide vehicles is the ability to attain the landing site substantially of the orbital plane and to make a horizontal landing. By widening both the entry window and the corridor the entry conditions become less stringent and operational flexibility is highly increased. The cross range capability depends on the L/D properties of the vehicle and the angle control, which simultaneously has to observe the flight path constraints. The maximum lateral range is a primary mission requirement and dictates vehicle angle.
The three factors influencing the baseline re-entry trajectory profile of winged vehicle are:
1. Temperature constraints, which primarily affect the vehicle, reuse and reflect specific TPS design.
2. Constraints to guarantee flight stability and system and payload integrity.
3. Down range and Cross range requirements to assure successful landing.

Space shuttle uses their wings to glide to runway and land like an airplane. For landing in non-atmospheric planet reverse thrust rockets are used by the vehicles. Figure given below explains the entry vehicle evolution.

Entry Vehicle evolution
In the design of a practical manned re-entry system, it is necessary to accomplish two things. The first is to bring the entry deceleration down to levels well within the human endurance, and the second is to limit the maximum heating rate and total heat pulse to less than material limits. A method of accomplishing these two goals leads to the following conclusions.
1. Re-entry at high total angle of attack provides a reduction in both maximum heating rate and total heating encountered.
2. Rolling maneuvers at constant total angle of attack are useful in reducing re-entry decelerations and total heating encountered.
3. A light wing loading combined with rolling entry maneuvers and a relatively blunt stagnation region permits a lightweight thermal protection system.
4. It is possible by suitable choice of configurations to obtain a vehicle with desirable aerodynamic characteristics during entry and good subsonic flying qualities for landing.
5. Extremely attractive payload to gross weight ratio of approximately 50% may be obtained in a vehicle of light wing loading if the aerodynamic and structural configurations are properly selected and rolling entry employed.
i. LOH, W.H.T: Re-entry and planetary entry physics and technology, VOL.2 & 3.
ii. Spacecraft Systems Engineering “ Peter Fortescue and John Stoaark.
iii. Hankey, W.L: Re-entry Aerodynamics, AIAA education series (American Institute of Aeronautics and Astronautics).
iv. Kojro and Takashi ABE: Viscous Shock-Layer Analysis on Hypersonic Flow over Re-entry Capsule.
v. P.N. Keert : Re-entry, Descent and Landing Technology and Demonstration needs for a European Winged Vehicle.

1. Introduction
2. Re-entry mission profile, constraints, and vehicle requirements
3. Entry corridor
4. Gas dynamics and deceleration
5. Aerodynamic heating
6. Material selection in design
7. Thermal protection system (TPS)
8. Vehicle guidance and landing
9. Conclusion
10. Reference

First of all I thank the almighty for providing me with the strength and courage to present the seminars.
I avail this opportunity to express my sincere gratitude towards
Dr. T.N. Sathyanesan, head of mechanical engineering department, for permitting me to conduct the seminars. I also at the outset thank and express my profound gratitude to my seminars guide Mr. Manjeeth Shukoor and staff incharge Asst. Prof. Mrs. Jumailath Beevi. D., for their inspiring assistance, encouragement and useful guidance.
I am also indebted to all the teaching and non- teaching staff of the department of mechanical engineering for their cooperation and suggestions, which is the spirit behind this report. Last but not the least, I wish to express my sincere thanks to all my friends for their goodwill and constructive ideas.


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