Nov, 2010

Soumik Bose Keshav Kishore Anand

Indian Institute Of Technology, Kharagpur

DEPARTMENT

OF

AEROSPACE

ENGINEERING

REUSABLE LAUNCH SYSTEMS

1

INTRODUCTION

A hundred years ago, on December 17, 1903, Wilbur and Orville Wright successfully achieved a

piloted, powered flight. Though the Wright Flyer I flew only 10 ft off the ground for 12 seconds,

traveling a mere 120 ft, the aeronautical technology it demonstrated paved the way for passenger

air transportation. Man had finally made it to the air. The Wright brother’s plane of 1903 led to

the development of aircrafts such as the WWII Spitfire, and others. In 1926 the first passenger

plane flew holiday makers from American mainland to Havana and Bahamas. In 23 years the

world had moved from a plane that flew 120 ft and similar planes that only a chosen few could

fly, to one that can carry many passengers.

In October 1957, man entered the space age. Russia sent the first satellite, the Sputnik, and in

April 1961, Yuri Gagarin became the first man on space. In the years since Russia and United

States has sent many air force pilots and a fewer scientists, engineers and others. But even after

almost 50 years, the number of people who has been to space is close to 500.

The people are losing interest in seeing a chosen few going to space and the budgets to space

research is diminishing. The space industry now makes money by taking satellites to space. But a

major factor here is the cost. At present to put a single kilogram into orbit will cost you between

$10000 and $20000. This is clearly prohibitively high and a major objective for the coming years

is to drop the cost to a fraction of today's value. Despite the fact that the space shuttle has

regularly gone into orbit over the last two decades there is still no tourist business. This is due to

the fact that to build an orbital hotel under present conditions will cost 100's of billions of dollars

(at least). It is clear why there has been so little progress in orbital developments.

The development of a plane which can fly to space at lower cost, which is reusable and can take

more payloads, is very much required for further development of space industries. The Reusable

Launch Vehicle, usually called Space plane or Hyper plane which can take crew and payload

into orbit is being developed by various space agencies and private companies. The Space plane

would make space travel cheap and will help in increasing space tourism and just like in the

aviation industry, within a few decades, the space tourism industries would be worth billions.

Reusable Launch System (or reusable launch vehicle, RLV)

It is a launch system which is capable of launching a launch vehicle into space more than once.

This contrasts with expendable launch systems, where each launch vehicle is launched once and

then discarded.

RLV's, due to the fact that they are re-used, will dramatically reduce the cost of access to low

earth orbit. However, the technical challenges of designing a system to fly to orbit and return are

monumental. For example, the entire Saturn V rocket was expended while sending humans to

the Moon. On the other hand, the Space Shuttle, which transports astronauts to Low Earth Orbit

and back, is reused over and over again.

A number of government research projects, most notably the X-33, X-34, X-37 and X-38, were

initiated to develop and test new RLV technologies to reduce the risk of developing a next

generation RLV. At the same time, a number of entrepreneurial companies have developed their

own RLV concepts in an effort to reduce launch costs and undercut established launch vehicle

providers.

Orbital RLVs are thought to provide the possibility of low cost and highly reliable access to

space. However, reusability implies weight penalties such as non-ablative reentry shielding and

possibly a stronger structure to survive multiple uses, and given the lack of experience with these

vehicles, the actual costs and reliability are yet to be seen.

No true orbital reusable launch system is currently in use. The closest example is the partially

reusable Space Shuttle. The orbiter, which includes the main engines, and the two solid rocket

boosters, are reused after several months of refitting work for each launch. The external fuel drop

tank is typically discarded, but it is possible for it be re-used in space for various applications.

Advantages

The rockets which take satellites and other payloads have to carry the fuel and oxidizer with

them as it uses conventional rocket engines. The combined weight of the fuel and oxidizer is

very large due to the fact that a lot of energy is expended pushing the plane forwards. This is

why today's rockets launch vertically as it maximizes the rocket's potential by allowing all the

energy expended to be focused in the direction we want to go-upwards. With present technology

it is the easiest and cheapest method of reaching space.

Clearly then the way forward is to utilize jet engines in some manner. The main advantages of jet

engines over rocket engines are that they do not need to carry their own oxidizer; instead they

suck in air and use the oxygen present in the air as their oxidizer. This will greatly remove the

need to carry oxidizer, as it will only be needed when at an altitude that the air contains

insufficient oxygen for jets to operate. At this point the rocket engines will fire and burn the

much smaller quantity of onboard oxidizer. This will dramatically reduce the take-off weight and

also the cost of the craft. Reduction in take-off weight means the payload can be increased.

Further to this the use of jet engines will make a substantial saving on the expensive rocket fuel.

As a comparison to produce the same thrust, jet (air-breathing) engines require less than one

seventh the propellants (fuel + oxidizer) that rockets do. For example, the space shuttle needs

143,000 gallons of liquid oxygen, which weighs 1,359,000 pounds (616,432 kg). Without the

liquid oxygen, the shuttle weighs a mere 165,000 pounds (74,842 kg).

Another advantage of jet-engine craft is that as they rely on aerodynamic forces rather than on

rocket thrust, they have greater maneuverability, which in turn provides better flexibility and

safety, for example missions can be aborted mid-flight if there is a problem. This is not the case

for staged vehicles, which typically have complex "range safety" requirements as the stages

detach and fall back to earth. Range safety is one of the main reasons that the US launches from

Florida, where the rocket's flight path takes it out over open water almost immediately. The lack

of such abort modes on the Shuttle requires incredible failure avoidance costs and massive

overhauls.

The space shuttle used by NASA is partially reusable. It still has to take off vertically with the

help of multistage rocket and solid boosters. The use of rockets increases the cost of

manufacturing parts for each launch as some rocket parts are not reusable. Furthermore, using

rockets increases the amount of fuel and oxidizer required. Some of the components of the rocket

get added to the space debris and continue orbiting the earth. This causes unwanted collisions

with other debris or satellites. Thus using a jet-engine craft as a reusable launch vehicle is faster,

efficient, and has increased affordability, flexibility and safety for ultra high-speed flights within

the atmosphere and into Earth orbit.

2

THE DIFFERENT REUSABILITY CONCEPTS

The difference is in way the RLV is launched. This difference results in differing levels of

complexity in design. The major concepts are:

Two Stage To Orbit (TSTO)

Two stage to orbit requires designing and building two independent vehicles and dealing with the

interactions between them at launch. Usually the second stage in launch vehicle is 5-10 times

smaller than the first stage, although in biamese and triamese approaches each vehicle is the

same size.

In addition, the first stage needs to be returned to the launch site for it to be reused. This is

usually proposed to be done by flying a compromise trajectory that keeps the first stage above or

close to the launch site at all times, or by using small airbreathing engines to fly the vehicle back,

or by recovering the first stage downrange and returning it some other way (often landing in the

sea, and returning it by ship.) Most techniques involve some performance penalty; these can

require the first stage to be several times larger for the same payload, although for recovery from

downrange these penalties may be small.

The second stage is normally returned after flying one or more orbits and reentering. An

advantage of such a system over single-stage-to-orbit is that the entire mass of the spacecraft is

not carried into orbit. This reduces the difficulty involved in reaching orbital velocity.

An advantage over three or more stages is reduction in complexity and fewer separation events,

each of which reduces cost and risk of failure.

One and a Half Stage To Orbit (OHSTO)

A classical rocket has one, two, three, four etc. stages to make orbit, each smaller than the

previous, each burning for a time, before the next one lights.

However most rockets, such as The Space Shuttle, have more than one set of engines running in

parallel at lift off. At some altitude one set (the half stage) of engines (and often associated fuel

tanks) are dropped off, and the remaining stage continues burning all the way to orbit. This is

known as One and a Half Stages to Orbit.

The biggest advantage of this idea is that total size and weight of the engines can be smaller and

lighter, as the 'half stage' engine gives lots of extra thrust early on, which means that more fuel

can be lifted, initially, and more payload.

In fact the half stage's engines give more thrust than the main ("sustainer") engines- so they quite

literally lob the rocket up. At separation the remaining stage actually doesn't have enough thrust

to lift the rocket on its own, not until it has burnt off most of the fuel; so the sustainer engines are

then in a race to burn off enough fuel so that the acceleration on the lighter rocket is sufficient to

make orbital velocity before falling back down to earth; although strangely enough, they usually

make it, by careful design.

Stage and a half designs have reliability advantages in that all engines are already lit at takeoff,

so there is less chance of a failure in the air- as most engine failures happen at ignition.

The disadvantage is that rocket nozzles work best when they deliver hot gas at the ambient

pressure- since as the rocket climbs the pressure changes, so the nozzle of the sustainer engine

has to be something of a compromise and loses some efficiency.

Single Stage To Orbit (SSTO)

It is a reusable launch vehicle (RLV) that takes off and lands horizontally like a conventional

plane. It is generally regarded that this method will be more efficient and safer than the 2-stage

model, though that is not to belittle the 2-stage method which would be a considerable

improvement on the vertical takeoff craft of today. It is also felt that while the 2-stage idea would

be easier, the 1-stage would almost certainly be more commercially viable and would achieve a

higher level of success in the objectives of a space plane.

What is required here is further development of jet engines. The only possibility at the moment is

ramjet working together with scramjet (Supersonic Combustion Ramjet). The major problem is

that the scramjets are far from fully developed, offering many difficult aerodynamic problems.

These, however, offer the only current hope of sustained hypersonic flight.

Even with the advance of scramjet development there are still many problems to be addressed

with horizontal take off of space planes. This is because a Scramjet will only function at

hypersonic speeds and a ramjet will only function at supersonic speeds.

The design will therefore require:

1. A turbojet, once the air intake reaches to mach 1 (supersonic speed) the ramjet would

fire.

2. The ramjet would accelerate the plane to about mach 4 (hypersonic speed) then the

scramjet would fire.

3. The scramjet is expected to be able to reach speeds of mach 15, when finally the rocket

engine would fire.

4. The rocket engine would accelerate the plane to mach 25 (escape velocity) and would be

used in space operations.

While this sounds very good in theory, in practice it is very doubtful whether such vehicles will

have the efficiency to reach orbit, due to the excessive weight and complexity of such a system.

Further to this such a design will not solve the other problem of heat build-up.

What we are really looking for is the development of a combined jet engine that operates across

the range, with maybe a switch to a rocket engine for the last stage and for space operations. The

difficulties of designing a jet engine to perform at these levels are such that it cannot even be

seen how it could be done with present technology. The differences between the engines are how

they physically take the air in. Nanotechnology could solve the problem by allowing the engine

to reshape itself in flight, whether it could be shaped fast enough remains to be seen.

Horizontal Landing

In this case the vehicle requires wings and undercarriage (unless landing at sea). This typically

requires about 9-12% of the landing vehicle to be wings; which in turn implies that the takeoff

weight is higher and/or the payload smaller.

Concepts such as lifting bodies attempt to deal with the somewhat conflicting issues of reentry,

hypersonic and subsonic flight; as does the delta wing shape of the Space Shuttle.

Fig: Horizontal Landing

Vertical Landing

Parachutes could be used to land vertically, either at sea, or with the use of small landing rockets,

on land (as with Soyuz).

Alternatively rockets could be used to softland the vehicle on the ground from the subsonic

speeds reached at low altitude (see DC-X). This typically requires about 10% of the landing

weight of the vehicle to be propellant.

A slightly different approach to vertical landing is to use an autogyro or helicopter rotor. This

requires perhaps 2-3% of the landing weight for the rotor.

Horizontal Take-Off

The vehicle needs wings to take off. For reaching orbit, a 'wet wing' would often need to be used

where the wing contains propellant. Around 9-12% of the vehicle takeoff weight is perhaps tied

up in the wings.

Vertical Take-Off

This is the traditional takeoff regime for pure rocket vehicles. Rockets are good for this regime,

since they have a very high thrust/weight ratio (~100).

Fig: Vertical Takeoff

Air-Breathing

Airbreathing approaches use the air for propulsion during ascent. The most commonly proposed

approach is the scramjet, but turborocket, Liquid Air Cycle Engine (LACE) and precooled

engines are also proposed to be used.

In all cases the highest speed that airbreathing can reach is far short of orbital speed (about Mach

15 for Scramjets and Mach 5-6 for the other engine designs) and rockets would be used for the

remaining 10-20 Mach for orbit.

The thermal situation for airbreathers (particularly scramjets) can be awkward; normal rockets

fly steep initial trajectories to avoid drag, whereas scramjets would deliberately fly through

relatively thick atmosphere at high speed generating enormous heating of the airframe. The

thermal situation for the other airbreathing approaches is much more benign, although is not

without its challenges.

3

WORKING

The working of a RLV can be divided into 4 stages

First stage – subsonic and supersonic stage

The RLV with its payload takes off from the runway and climbs to about 100,000 feet or 30km

using conventional jet-engines, or using a combination of conventional jet-engine and ramjet

engine, or using another plane to carry or pull the plane to a lower height and using a booster

rocket.

A ramjet operates by subsonic combustion of fuel in a stream of air compressed by the forward

speed of the aircraft. It doesn’t have or use very less moving parts compared to a conventional

jet-engine with thousands of moving parts. The compression of air before burning of fuel is done

in the ramjet by the addition of a diffuser at the inlet, while it is done by the turbine in

conventional jet-engine. The flow of air is subsonic.

The plane is accelerated to a speed of mach 4 or mach 5 and the flow inside the engine becomes

supersonic. Then the scramjet is powered up.

Second stage - Hypersonic stage:

When the space plane is at an altitude of about 100,000 ft and at a velocity of about mach 4, the

scramjets are fired.

Scramjets are basically ramjets. They introduce fuel and mix it with oxygen obtained from the air

which compressed for combustion. The air is naturally compressed by the forward speed of the

vehicle and the shape of the inlet, similar to what turbines or pistons do in slower-moving

airplanes and cars. Rather than using a rotating compressor, like a turbojet engine does, the

forward velocity and aerodynamics compress the air into the engine. Hydrogen fuel is then

injected into the air stream, and the expanding hot gases from combustion accelerate the exhaust

air to create tremendous thrust. While the concept is simple, proving the concept has not been

simple. At operational speeds, flow through the scramjet engine is supersonic - or faster than the

speed of sound. At that speed, ignition and combustion take place in a matter of milliseconds.

This is one reason it has taken researchers decades to demonstrate scramjet technologies, first in

wind tunnels and computer simulations, and only recently in experimental flight tests.

The Scramjet engine takes the RLV to even greater heights and to speeds of up to Mach 15. This

is the fastest speed an air breathing plane can go using current technologies. At Mach 15, the

RLV is at a great height that there isn’t enough oxygen to sustain the scramjet engine. At this

point the rocket engine fires up.

Third stage - Space stage:

When the rocket engine fires by mixing oxygen from the onboard storage tanks into the scramjet

engine, thereby replacing the supersonic airflow. The rocket engine is capable of accelerating the

RLV to speeds of about Mach 25, which is the escape velocity. It takes the RLV into orbit. The

rocket engine takes the RLV to the payload release site and the required operations are done.

Once this is over it enters its last stage – the re-entry stage.

Fourth stage – Re-entry stage:

Once the RLV finishes its mission in space, It performs de-orbit operations, including firing its

thrusters to slow itself down, thereby dropping to a lower orbit and eventually entering the upper

layers of the atmosphere. As the vehicle encounters denser air, the temperature of the ceramic

skin builds to over 1,000 degrees C, and is also cooled by using any remaining liquid hydrogen

fuel. It is here that the structure of the plane undergoes heavy thermal stress. If the heat shields

do not protect the plane, it would simply burn off to the ground, just like the space shuttle

Columbia. It enters a radio silence zone as due to the heat, radio contact is lost. Once it reaches

dense air, it can use its aerodynamics to glide down to the landing strip. It can also use any

remaining fuel to fire the ramjet or conventional jet (depends on the design) and change its

course. Once on the landing strip it engages it slows down using a series of parachutes and

engages the brake.

4

CONSTRUCTION AND DESIGN

The construction of a true RLV that can take a payload to space is still in the design stage. It will

sure have lot of designs taken from the space shuttle. Here are the main construction details.

Body

The body of a RLV has to withstand very high stresses including thermal stresses during re-

entry. The plane expands due to the high heat of nearly 1500˚C or more. It also has to cope with

the rapid change in temperatures once in space. It changes from -250 degrees in the shade to 250

degrees in direct sunlight. This change in temperature between two sides of the same plane will

put a lot of stress on its body. Titanium alloys are being used, being very strong and light. To

cope with the high temperatures developed in parts of the wing and fuselage of the spacecraft

today, reinforced carbon-carbon composite material is being added to the leading edges of the

vehicle's nose and wings to handle the higher temperatures.

Researchers are being conducted to find the best materials for different parts of the plane. One of

these materials, γ-TiAl (Titanium Aluminide), has superior high-temperature material properties.

Its low density provides improved specific strength and creep resistance in comparison to

currently used titanium alloys. However, it is inherently brittle, and long life durability is a

potential problem along with the material’s sensitivity to defects.

Wings

The wing of the spacecraft has to be designed so that it provides enough lift to fly to space and

also reduce the friction during re-entry.

Cockpit

The cockpit is the place where the astronauts will stay most of the time during the journey. It will

have many windows, which are special double-paned glass, and each pane alone can withstand

the pressure and force of flight and the vacuum. This doubling up ensures that if either window

were to crack, the passengers would still be safe.

The air inside the cockpit is made breathable by a three-part system. Breathable air is added at a

constant rate by oxygen bottles. The exhaled carbon dioxide is removed from the cabin by an

absorber system, and humidity is controlled by an additional absorber created to remove water

vapor from the air. During the entire flight, the cockpit remains comfortable, cool and dry.

The avionics system and display unit for navigating has to be computer controlled and free from

bugs. It should give the pilot all the necessary data to make his choices. The avionics are very

critical, and it also needs to be very precise for the pilot to do what he wants to do, and do it well.

Electric Power

The electrical power required for the running of the spacecraft has to be taken from batteries.

These batteries could be charged, if needed by using solar energy. Researches are being initiated

to find better and reliable batteries, like the lithium-based (i.e., Li metal or Li-ion intercalation

compound as negative electrode), polymer electrolyte regenerative battery system. Its advantages

include reduced battery weight and volume, relative to conventional Ni-Cd and Ni-H2, which

permits greater payloads and greater cell voltage, 3.5 volts vs. 1.2 volts, which permits use of

fewer cells and results in reduced battery system complexity.

Controls

When we're out in space, all you need to do is release a puff of air in a direction to give you a

reaction force to push you the other way. That is called a reaction control system. High-pressure

air is stored in bottles on the ship, and on the release of a little blast of air for about one second,

for example, with the right wing tip pointing up. And that is enough when you're in space to push

that wingtip down. It effectively rolls the aircraft, and that are the controls when it is out in

space.

Fuels

Many challenges have been overcome recently by the discovery and synthesis of propellants that

can have higher performance than traditional O2/H2, and aircraft fuels. These propellants include

high-density monopropellants for sounding rockets and upper stages, and onboard propulsion for

small spacecraft. Higher energy fuels, such as N4, N6, BH4, and others, have a longer range

development time and would be more applicable to future launch vehicles.

Heat Shields

Reentry heat shields on these vehicles are often proposed to be some sort of ceramic and/or

carbon-carbon heat shields, or occasionally metallic heat shields (possibly using water cooling or

some sort of relatively exotic rare earth metal.)

Some shields would be single use ablatives and would be discarded after reentry.

A newer Thermal Protection System (TPS) technology was first developed for use in steering

fins on ICBM MIRVs. Given the need for such warheads to reenter the atmosphere swiftly and

retain hypersonic velocities to sea level, researchers developed what are known as SHARP

materials, typically hafnium diboride and zirconium diboride, whose thermal tolerance exceeds

3600 C. SHARP equipped vehicles can fly at Mach 11 at 30 km altitude and Mach 7 at sea level.

The sharp-edged geometries permitted with these materials also eliminates plasma shock wave

interference in radio communications during reentry. SHARP materials are very robust and

would not require constant maintenance, as is the case with technologies like silica tiles, used on

the Space Shuttle, which account for over half of that vehicles maintenance costs and turnaround

time. The maintenance savings alone are thus a major factor in favor of using these materials for

a reusable launch vehicle, whose raison d'etre is high flight rates for economical launch costs.

Thermal Protection Systems (TPS)

Ablative

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) through blowing providing the best protection

against high heat flux. The overall process of reducing the heat flux experienced by the heat

shield's outer wall is called blockage. Ablation causes the TPS layer to char, melt, and sublime

through the process of pyrolysis. 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. Ablation

can also provide blockage against radiative heat flux by introducing carbon into the shock layer

thus making it optically opaque. 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 RV nose tips. Thermal protection can also be enhanced in some TPS materials through

coking.

Early research on ablation technology in the USA was centered at NASA's Ames Research

Center located at Moffett Field, California. 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

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

phenolic is a very effective ablative material, but also has high density which is undesirable. 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:

i) SLA-561V

SLA" in SLA-561V stands for "Super Light weight Ablator". SLA-561V is a proprietary ablative

made by Lockheed Martin that has been used as the primary TPS material on all of the 70 degree

sphere-cone entry vehicles sent by NASA to Mars. SLA-561V begins significant ablation at a

heat flux of approximately 110 W/cm², but will fail for heat fluxes greater than 300 W/cm². SLA-

561V 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.

ii) PICA

Phenolic Impregnated Carbon Ablator (PICA) was developed by NASA Ames Research Center

and was the primary TPS material for the Stardust aeroshell. The Stardust sample-return capsule

was the fastest man-made object ever to reenter Earth's atmosphere (12.4 km/s or 28,000 mph at

135 km altitude) This was faster than the Apollo mission capsules and 70% faster than the

Shuttle. PICA was critical for the viability of the Stardust mission.

PICA is a modern TPS material and has the advantages of low density (much lighter than carbon

phenolic) coupled with efficient ablative capability at high heat flux. Stardust's heat shield (0.81

m base diameter) was manufactured from a single monolithic piece sized to withstand a nominal

peak heating rate of 1200 W/cm2. PICA 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.

iii) SIRCA

Silicone Impregnated Reusable Ceramic Ablator (SIRCA) was also developed at NASA Ames

Research Center. SIRCA is a monolithic, insulative 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 current 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. SIRCA has been demonstrated in BIP applications, but not yet as a

forebody TPS material.

Thermal Soak

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. This outcome is prevented by ejecting the heat shield (with its heat soak) prior to the

heat conducting to the inner wall.

Thermal soak TPS is intended to shield mainly against heat load and not against a high peak heat

flux (a long duration heat pulse of low intensity is assumed for the TPS design). The Space

Shuttle orbit vehicle was designed with a reusable heat shield based upon a thermal soak TPS. It

should be emphasized that the tradeoff for TPS reusability is an inability to withstand a high heat

flux; e.g., a Space Shuttle TPS would not be practical as a primary thermal protection for lunar

return. A Space Shuttle's underside is coated with thousands of tiles made of silica foam, which

are intended to survive multiple reentries with only minor repairs between missions. Fabric

sheets known as gap fillers are inserted between the tiles where necessary. These gap fillers

provide for a snug fit between separate tiles while allowing for thermal expansion. When a Space

Shuttle lands, a significant amount of heat is stored in the TPS. Shortly after landing, a ground-

support cooling unit connects to the Space Shuttle's internal Freon coolant loop to remove heat

soaked in the TPS and orbiter structure.

Typical Space Shuttle's TPS tiles (LI-900) have remarkable thermal protection properties, but are

relatively brittle and break easily, and cannot survive in-flight rain. An LI-900 tile exposed to a

temperature of 1000 K on one side will remain merely warm to the touch on the other side. An

impressive stunt that can be performed with a cube of LI-900 is to remove it glowing white hot

from a furnace and then hold it with one's bare fingers without discomfort along the cube's edges.

Fig: Silica tile material, safely hand held 10 seconds after removal from high temperature oven

Passively Cooled

In some early ballistic missile RVs, e.g., the Mk-2 and the sub-orbital Mercury spacecraft,

radiatively cooled TPS were used to initially absorb heat flux during the heat pulse and then after

the heat pulse, radiate and convect the stored heat back into the atmosphere. However, the earlier

version of this technique required a considerable quantity of metal TPS (e.g., titanium, beryllium,

copper, et cetera). Modern designers prefer to avoid this added mass by using ablative and

thermal soak TPS instead.

Radiatively cooled TPS can still be found on modern entry vehicles, but Reinforced Carbon-

Carbon (also called RCC or carbon-carbon) is normally used instead of metal. RCC is the TPS

material on the nose cone and leading edges of the Space Shuttle's wings. RCC was also

proposed as the leading edge material for the X-33. Carbon is the most refractory material known

with a one atmosphere sublimation temperature of 3825 °C for graphite. This high temperature

made carbon an obvious choice as a radiatively cooled TPS material. Disadvantages of RCC are

that it is currently very expensive to manufacture and lacks impact resistance.

A radiatively cooled TPS for an entry vehicle is often called a "hot metal TPS". Early TPS

designs for the Space Shuttle called for a hot metal TPS based upon nickel superalloy (Rene-41)

and titanium shingles the earlier Shuttle TPS concept was rejected because it was incorrectly

believed a silica tile based TPS offered less expensive development and manufacturing costs. A

nickel superalloy shingle TPS was again proposed for the unsuccessful X-33 Single-Stage to

Orbit (SSTO) prototype.

Actively Cooled

Various advanced reusable spacecraft and hypersonic aircraft designs have been proposed to

employ heat shields made from temperature-resistant metal alloys that incorporated a refrigerant

or cryogenic fuel circulating through them. Such a TPS concept was proposed for the X-30

National Aerospace Plane (NASP). The NASP was supposed to have been a scramjet powered

hypersonic aircraft, but failed in development.

In the early 1960s 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, more rugged, and

eliminating the need for classified technology. The disadvantage is increased weight and

complexity, and lower reliability. The concept has never been flown, but a similar technology

did undergo extensive ground testing.

Re-entry Module Designs

Sphere-cone

The sphere-cone is a spherical section with a frustum or blunted cone attached. The sphere-

cone's dynamic stability is typically better than that of a spherical section. 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.)

Fig: Sphere-cone arrangement\

Reconnaissance satellite RVs (recovery vehicles) also used a sphere-cone shape and were the

first American example of a non-munition entry vehicle. The sphere-cone was later used for

space exploration missions to other celestial bodies or for return from open space; e.g., Stardust

probe. Unlike with military RVs, the advantage of the blunt body's lower TPS mass remained

with space exploration entry vehicles like the Galileo Probe with a half angle of 45° or the

Viking aeroshell with a half angle of 70°. Space exploration sphere-cone entry vehicles have

landed on the surface or entered the atmospheres of Mars, Venus, Jupiter and Titan.

Biconic

The biconic is a sphere-cone with an additional frustum attached. The biconic offers a

significantly improved L/D ratio. The higher L/D makes a biconic shape better suited for

transporting people to Mars due to the lower peak deceleration. Arguably, the most significant

biconic ever flown was the Advanced Maneuverable Reentry Vehicle (AMaRV). Four AMaRVs

were made by the McDonnell-Douglas Corp. and represented a significant leap in RV

sophistication.

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. The McDonnell Douglas DC-X (also a

biconic) was essentially a scaled up version of AMaRV. AMaRV and the DC-X also served as

the basis for an unsuccessful proposal for what eventually became the Lockheed Martin X-33.

Amongst aerospace engineers, AMaRV has achieved legendary status alongside such marvels as

the SR-71 Blackbird and the Saturn V Rocket.

Non-axisymmetric shapes

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. This approach has been used by the American Space Shuttle and the Soviet Buran. The

lifting body is another entry vehicle geometry and was used with the X-23 PRIME (Precision

Recovery Including Maneuvering Entry) vehicle.

The FIRST (Fabrication of Inflatable Re-entry Structures for Test) system was an Aerojet

proposal for an inflated-spar Rogallo wing made up from Inconel wire cloth impregnated with

silicone rubber and Silicon Carbide dust. FIRST was proposed in both one-man and six man

versions, used for emergency escape and reentry of stranded space station crews, and was based

on an earlier unmanned test program that resulted in a partially successful reentry flight from

space (the launcher nose cone fairing hung up on the material, dragging it too low and fast for

the TPS, but otherwise it appears the concept would have worked, even with the fairing dragging

it, the test article flew stably on reentry until burn-through).

Feathered Re-entry

In 2004, aircraft designer Burt Rutan demonstrated the feasibility of a shape changing airfoil for

reentry with the suborbital SpaceShipOne. The wings on this craft rotate to provide a shuttlecock

effect. Notably, SpaceShipOne does not experience significant thermal loads on reentry.

This increases drag, as the craft is now less streamlined. This 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 very key to efficient reentry). It should also be

noted that SpaceShipOne, in its "wings flipped" configuration, will automatically orient itself to

a high drag attitude. Rutan has compared this to a falling shuttlecock. However, it is important to

realize that the velocity attained by SpaceShipOne prior to reentry is much lower than that of an

orbital spacecraft, and most engineers do not consider the shuttlecock reentry technique viable

for return from orbit.

Entry Vehicle Design Considerations

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

Peak heat flux

Heat load

Peak deceleration

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. 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-averse designer).

Maximum aeroshell bluntness (maximum drag) yields minimum TPS mass. Maximum bluntness

(minimum ballistic coefficient) also yields a minimal terminal velocity at maximum altitude.

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.

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.

5

CONCLUSION

The recent development in the field of scramjets, hyperplanes, and truly Reusable Launch

Vehicle will result in the development of space tourism.

A list of different reusable launch vehicles is given in the list below:

Currently in use:

Space Shuttle - Partially reusable

Planned:

PlanetSpace Silver Dart - Partly reusable spaceplane, based on hypersonic glider design.

SpaceX Falcon 1 - Announced as partially reusable; 28 September 2008 test flight reached orbit,

but vehicle recovery not yet demonstrated.

SpaceX Falcon 9 - Announced as partially reusable; maiden flight achieved orbit, 4 June 2010.

Reaction Engines Skylon - Proposed airbreathing SSTO spaceplane.

Avatar RLV - Proposed reusable Indian launch system for small payloads

Historical:

Soviet Union Energia - Buran system (partially reusable, now cancelled)

In 1927, hotel magnate Raymond Orteig's announced an aviation contest, a price of $25,000 to

be awarded to the first man to build and fly an airplane non-stop from New York to Paris. As a

result of the successful flight of Charles Lindbergh's, in the United States:

The number of airline passengers increased by 167,623 between 1926 and 1929.

The number of pilot's license applications increased by 300 percent in 1927.

The number of licensed aircraft increased by 400 percent.

The number of airports doubled within three years.

Once this technology will be common, people will start their pursuit towards even better

technologies. Only time will tell if the Ansari X Prize will have a similar effect on the

burgeoning sub-orbital flight industry as the 1927 $25,000 Orteig's price did.

6

BIBLIOGRAPHY

www.nasa.gov

www.spaceandtech.com

www.hobbyspace.com

www.wikipedia.org

www.everything2.com