REENTRY MECHANISMS TO REDUCE THE HEAT LOADS ON SPACE CRAFTS

By

T.Sugavanan (07AE71)

G.Vasanth Kumar (07AE74)

Abstract

Atmospheric reentry is the movement of human-made or natural objects as they enter the atmosphere of a planet from outer space, in the case of Earth from an altitude above the Kármán Line, (100 km). Vehicles that typically undergo this process include ones returning from orbit (spacecraft) and ones on exo-orbital (suborbital) trajectories (ICBM reentry vehicles, some spacecraft.). Typically this process requires special methods to protect against aerodynamic heating. Various advanced technologies have been developed to enable atmospheric reentry and flight at extreme velocities. For different air borne vehicles to enter into the atmosphere of heavenly bodies, different entry mechanisms are adopted

This paper is about such technologies and mechanisms which seemed far-fetched decades earlier. To list a few we will briefly explain the skip reentry mechanism, double dip reentry, feathered reentry.

REENTRY MECHANISMS TO REDUCE THE HEAT LOADS ON SPACE CRAFTS

INTRODUCTION

When one thinks of atmospheric re-entry, they commonly think of such things as space shuttles coming back from a mission. However, while the term may be slightly erroneous is some cases, it also involves satellites and meteors that burn up in our atmosphere as well as those that manage to land. We, however, are mostly concerned with the satellites as they have little to no propulsion system of their own and originated from the earth’s surface. Whether they burn up or not is a matter of thermodynamics, but where they land is a matter of mechanics.

The concept of the ablative heat shield was described as early as 1920 by Robert Goddard. Practical development of reentry systems began as the range and reentry velocity of ballistic missiles increased. For early short-range missiles, like the V-2, stabilization and aerodynamic stress were important issues (many V-2s broke apart during reentry), but heating was not a serious problem. Medium-range missiles like the Soviet R-5, with a 1200 km range, required ceramic composite heat shielding on separable reentry vehicles (it was no longer possible for the entire rocket structure to survive reentry). The first ICBMs, with ranges of 8000 to 12,000 km, were only possible with the development of modern ablative heat shields and blunt-shaped vehicles. In the USA, this technology was pioneered by H. Julian Allen at Ames Research. In the Soviet Union, Yuri A. Dunaev developed similar technology at the Leningrad Physical-Technical Institute.

CHALLENGES ENCOUNTERED

Before the disintegration of OV-102 Columbia highlighted the dangers of atmospheric re-entry, it had seemingly become almost routine; the Columbia Accident Investigation Board noted that NASA had a zero impact specification even though it recorded a multitude of thermal protection system (TPS) strikes by debris in every mission. Unfortunately, a piece of foam insulation fell off the External Tank and impacted the reinforced carbon-carbon leading edge of Columbia's wing, allowing the heat of re-entry to penetrate into the wing and cause it to fail. Since we had never suffered a catastrophic failure in this environment, the situation was assumed to be acceptable. Now, at great cost, we know better.

Dissipating kinetic energy

For any given orbital altitude, there is a specific velocity required to maintain that orbit. For a standard Space Shuttle LEO (Low-Earth Orbit) profile, the orbital velocity is around 17,500 mph (about 7823 m/s). The landing speed of the Shuttle is around 225 mph or 100 m/s (this velocity is even lower for reentry capsules, for obvious reasons). Given that kinetic energy equals ½*m*v^2, that equates to a factor of 60 difference between the two energy states. Somehow, a reentry vehicle needs to bleed off that excess energy. Enough fuel could be carried for a fully powered reentry, but that would be prohibitive from a mass standpoint and would still require you to kill all of your forward velocity to land. Given our available options, aerobraking, using the atmosphere to slow down was the natural choice. (The Delta Clipper, or DC-X, had it been built, would have been an exception to the rule, since it was to land vertically.)

Aerodynamic effects

From an aerodynamic standpoint, a vehicle can be blunt (a ballistic reentry vehicle, or R/V, like the Apollo capsule) or generate lift (like the shuttle). The first benefits from the strong shockwave it generates - this insulates it from the high-temperature airstream to a degree. The corresponding disadvantage is that it decelerates much more strongly, especially in the lower, denser parts of the atmosphere, though it requires less time to reenter. The lifting body slows down much more gradually by shifting its deceleration into the upper atmosphere.

Thermal effects

With either approach, you have to deal with the thermal consequences of your choice. Since the ballistic R/V gets to the ground faster, it’s also bleeding off kinetic energy faster, which results in correspondingly higher peak temperatures. The lifting R/V spends more time bleeding energy off in the upper atmosphere, so its peak heating is potentially lower. However, the total amount of heat absorbed can end up being higher. Furthermore, peak temperature and the sharpness of a surface are related, since the sharp edge doesn’t build up that insulating buffer of air that the ballistic R/V does. As a result, temperatures along the nose and leading edge of the lifting R/V can become very high (up to 2400 degrees F in the case of the Shuttle, and neither its leading edge nor its nose are particularly sharp). Constructing materials and structures that can withstand such temperatures is quite a challenge; making them durable and impact-resistant is an even greater challenge.

Entry vehicle design considerations

There are four critical parameters considered when designing a vehicle for atmospheric entry:

1. Peak heat flux2. Heat load3. Peak deceleration4. Peak dynamic pressure

Peak heat flux and dynamic pressure selects the TPS material. Heat load selects the thickness of the TPS material stack. Peak deceleration is of major importance for manned missions. The upper limit for manned return to Earth from Low Earth Orbit (LEO) or lunar return is 10 Gs. For Martian atmospheric entry after long exposure to zero gravity, the upper limit is 4 Gs. Peak dynamic pressure can also influence the selection of the outermost TPS material if spallation is an issue.

Starting from the principle of conservative design, the engineer typically considers two worst case trajectories, the undershoot and overshoot trajectories. The undershoot trajectory is typically defined as the shallowest allowable entry velocity angle prior to atmospheric skip-off. The overshoot trajectory has the highest heat load and sets the TPS thickness. The undershoot trajectory is defined by the steepest allowable trajectory. For manned missions the steepest entry angle is limited by the peak deceleration. The undershoot trajectory also has the highest peak heat flux and dynamic pressure. Consequently the undershoot trajectory is the basis for selecting the TPS material. There is no "one size fits all" TPS material. A TPS material that is ideal for high heat flux may be too conductive (too dense) for a long duration heat load. A low density TPS material might lack the tensile strength to resist spallation if the dynamic pressure is too high. A TPS material can perform well for a specific peak heat flux, but fail catastrophically for the same peak heat flux if the wall pressure is significantly increased (this happened with NASA's R-4 test spacecraft). Older TPS materials tend to be more labor intensive and expensive to manufacture compared to modern materials. However, modern TPS materials often lack the flight history of the older materials (an important consideration for a risk adverse designer).

Based upon Allen and Eggers discovery, maximum aeroshell bluntness (maximum drag) yields minimum TPS mass. Maximum bluntness (minimum ballistic coefficient) also yields a minimal terminal velocity at maximum altitude (very important for Mars EDL, but detrimental for military RVs). However, there is an upper limit to bluntness imposed by aerodynamic stability considerations based upon shock wave detachment. A shock wave will remain attached to the tip of a sharp cone if the cone's half-angle is below a critical value. This critical half-angle can be estimated using perfect gas theory (this specific aerodynamic instability occurs below hypersonic speeds). For a nitrogen atmosphere (Earth or Titan), the maximum allowed half-angle is approximately 60°. For a carbon dioxide atmosphere (Mars or Venus), the maximum allowed half-angle is approximately 70°. After shock wave detachment, an entry vehicle must carry significantly more shocklayer gas around the leading edge stagnation point (the subsonic cap).

Consequently, the aerodynamic center moves upstream thus causing aerodynamic instability. It is incorrect to reapply an aeroshell design intended for Titan entry (Huygens probe in a nitrogen atmosphere) for Mars entry (Beagle-2 in a carbon dioxide atmosphere). Prior to being abandoned, the Soviet Mars lander program achieved no successful landings (no useful data returned) after multiple attempts. The Soviet Mars landers were based upon a 60° half-angle aeroshell design. In the early 1960s, it was incorrectly believed the Martian atmosphere was mostly nitrogen, (actual Martian atmospheric mole fractions are carbon dioxide 0.9550, nitrogen 0.0270 and argon 0.0160).

A 45 degree half-angle sphere-cone is typically used for atmospheric probes (surface landing not intended) even though TPS mass is not minimized. The rationale for a 45° half-angle is to have either aerodynamic stability from entry-to-impact (the heat shield is not jettisoned) or a short-and-sharp heat pulse followed by prompt heat shield jettison. A 45° sphere-cone design was used with the DS/2 Mars impactor and Pioneer Venus Probes.

BLUNT BODY ENTRY VEHICLES

In the United States, H. Julian Allen and A. J. Eggers, Jr. of the National Advisory Committee for Aeronautics (NACA) made the counterintuitive discovery in 1951that a blunt shape (high drag) made the most effective heat shield. From simple engineering principles, Allen and Eggers showed that the heat load experienced by an entry vehicle was inversely proportional to the drag coefficient, i.e. the greater the drag, the less the heat load. Through making the reentry vehicle blunt, air can't "get out of the way" quickly enough, and acts as an air cushion to push the shock wave and heated shock layer forward (away from the vehicle). Since most of the hot gases are no longer in direct contact with the vehicle, the heat energy would stay in the shocked gas and simply move around the vehicle to later dissipate into the atmosphere.

Figure 1 Prototype Version of the Mk-2 Reentry Vehicle (RV) was derived from Blunt Body Theory

The Allen and Eggers discovery, though initially treated as a military secret, was eventually published in 1958.[4] The Blunt Body Theory made possible the heat shield designs that were embodied in the Mercury, Gemini and Apollo space capsules, enabling astronauts to survive the fiery reentry into Earth's atmosphere.

In the Soviet Union the R-7 ICBM was first successfully tested in 1957 with a sharp-nosed conical warhead. This burned up at an altitude of 10 km over the target area, and was replaced with a blunt-nosed conical design. Soviet heat shields consisted of layers of fiberglass together with asbestos textolite.

DIFFERENT REENTRY MECHANISMS

SKIP REENTRY

Skip reentry is a reentry technique involving one or more successive "skips" off the atmosphere to achieve greater entry range or to slow the spacecraft before final entry, which helps to dissipate the huge amount of heat that is usually generated on faster descents. The range modulation made possible by skip entry allows a spacecraft to reach a wider landing area, or to reach a designated landing point from a wider range of possible entry times, which is especially important in abort situations. Like aerocapture, skip reentry requires precise guidance. An overly shallow entry angle will result in the spacecraft retaining too much of its velocity, possibly escaping into space permanently if this is more than escape velocity. An overly steep entry, on the other hand, results in more intense heating and stress that could exceed the design limits of the spacecraft, potentially destroying it.

The basic concept is to 'clip' the atmosphere at such an angle that the craft is 'pushed' back out into space, conceptually similar to a pebble skipping across the surface of a lake. Each time, the craft's velocity is reduced so that it can eventually drop into the atmosphere at a low suborbital velocity.

In theory, any craft could perform skip re-entry as it does not require much lift, but in practise it requires precise guidance. Although the space shuttle is capable of skip reentry, NASA has carried it out only in computer simulations (Scott Horowitz, NASA interview, Jan. 25, 2007). It is unclear how thermal shielding would fare under the rapid heating, cooling and reheating. In theory STS-107 might have survived if a skip trajectory had been attempted - giving more time for heat dissipation - but this cannot be proven.

Figure 2Atmospheric entry trajectory illustrating the basic phases of flight in a skip reentry

Skip entry was first imagined in the 1930s, when a suborbital skipping trajectory was planned for the German Silbervogel bomber, which never flew. The technique was used by the Zond series of circumlunar spacecraft, which planned for one skip before landing. Zond 6 and 7 made successful skip entries, although Zond 6 was largely destroyed on impact for unrelated reasons. The Apollo Command Module, when returning from the moon, was capable of a one-skip entry. The Orion spacecraft crew module will be capable of skip entry, to allow targeting the landing site from a greater variety of abort trajectories.

DOUBLE DIP REENTRY

Double dip entry, takes advantage of the lift of the vehicle to make two decelerating dips into the atmosphere

the heat transferred is more gradual, allowing the use of a thinner heat shield that relies on radiating the heat

SPACECRAFT MISSIONS

1969 August 7 - Zond 7 - Program: Lunar L1. Launch Site: Baikonur. Launch Complex: LC81/23. Launch Vehicle: Proton. Mass: 5,379 kg (11,858 lb). Perigee: 200 km (120 mi). Apogee: 400,000 km (240,000 mi).

Inclination: 51.50 deg. Period: 15,562.22 min. Duration: 5.00 days.

Circumlunar flight; successfully recovered in USSR August 13, 1969. Only completely successful L1 flight that could have returned cosmonauts alive or uninjured to earth. Official mission was further studies of

the moon and circumlunar space, to obtain color photography of the earth and the moon from varying distances, and to flight test the spacecraft systems. Earth photos were obtained on August 9, 1969. On August 11, 1969, the spacecraft flew past the moon at a distance of 1984.6 km and conducted two picture taking sessions. Successfully accomplished double-dip re-entry and landed 50 km from aim point near Kustani in the USSR.

AEROBRAKING

Aerobraking is a spaceflight maneuver that reduces the high point of an elliptical orbit (apoapsis) by flying the vehicle through the atmosphere at the low point of the orbit (periapsis). The resulting drag slows the spacecraft. Aerobraking is used when a spacecraft requires a low orbit after arriving at a body with an atmosphere, and it requires less fuel than does the direct use of a rocket engine.

Method

When an interplanetary vehicle arrives at its destination, it must change its velocity to remain in the vicinity of that body. When a low, near-circular orbit around a body with substantial gravity (as is required for many scientific studies) is needed, the total required velocity changes can be on the order of several kilometers per second. If done by direct propulsion, the rocket equation dictates that a large fraction of the spacecraft mass must be fuel. This in turn means the spacecraft is limited to a relatively small science payload and/or the use of a very large and expensive launcher. Provided the target body has an atmosphere, aerobraking can be used to

reduce fuel requirements. The use of a relatively small burn allows the spacecraft to be captured into a very elongated elliptic orbit. Aerobraking is then used to circularize the orbit. If the atmosphere is thick enough, a single pass through it can be sufficient to slow a spacecraft as needed. However, aerobraking is typically done with many orbital passes through a higher altitude, and therefore thinner, region of the atmosphere. This is done to reduce the effect of frictional heating, and because unpredictable turbulence effects, atmospheric composition, and temperature make it difficult to accurately predict the decrease in speed that will result from any single pass. When aerobraking is done in this way, there is sufficient time after each pass to measure the change in velocity and make any necessary corrections for the next pass. Achieving the final orbit using this method takes a long time (e.g., over six months when arriving at Mars), and may require several hundred passes through the atmosphere of the planet or moon. After the last aerobraking pass, the spacecraft must be given more kinetic energy via rocket engines in order to raise the periapsis above the atmosphere--unless, of course, the intent is to land the spacecraft.

The kinetic energy dissipated by aerobraking is converted to heat, meaning that a spacecraft using the technique needs to be capable of dissipating this heat. The spacecraft must also have sufficient surface area and structural strength to produce and survive the required drag, but the temperatures and pressures associated with aerobraking are not as severe as those of reentry or aerocapture. Simulations of the Mars Reconnaissance Orbiter aerobraking use a force limit of 0.35 N per square meter with a spacecraft cross section of about 37 m², and a maximum expected temperature as 340 °F (170 °C). The force density, of roughly 0.2 N (0.04 lbf) per square meter[2], that was exerted on the Mars Observer, during aerobraking is comparable to the force of a 40 mph (60 km/h) wind on a human hand at sea level on Earth.

CONCLUSION

Now that we have taken a look at the various factors that can affect an object during atmospheric re-entry, we are left with, in the general case of an ellipse that intersects with the Earth’s surface, a general deformation of said ellipse by the pressure gradient force.

As for the many other factors that one might think would affect the trajectory of an object, they generally either have a negligible effect or only affect the path of the object while it is still in “orbit.” While they could possibly have an effect on one’s calculations of the path of an object based on phase 3 of a trajectory, they will not have an appreciable effect on a heavy object during phases 1 or 2.

The total effect that all of this has on the trajectory is a shallowing of the path on launch and a steepening of the path upon reentry. Thus, this is the reason that most objects made for re-

entry fulfill have their own propulsion system or are designed to re-enter the atmosphere at a low angle to begin with. Most designs utilize both so as to conserve fuel. Once the space capsule, for example, has slowed down sufficiently they then make use of parachutes to further slow the descent so as to make for a softer landing.