The document contains the following sections:

Section A: Basics of Space Transport

Section B: Propulsive Forces List

Section C: Energy Sources List

Section D: Propulsion Concepts List

Section E: Space Engineering Methods

Section F: General References

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0.1-0.7 <5/94
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The traditional job of the rocket designer has been to find the best compromise between high cost and small payload when going to Earth orbit. Larger payloads can be achieved by making a rocket last a single flight (thus using lighter structures than ones built to last many flights), and by dropping parts of the propulsion system (as fuel is used up less thrust is required to maintain acceleration, so you can drop engines). These measures are expensive (you have to replace or re-assemble the rocket), but were necessary in the past because of the weight of structures and the low performance of chemical rockets.

Today the job of the space transportation system designer is much more complex. There are many more propulsion concepts available (over 60 in this document alone), and missions are not limited to getting into Earth orbit. Technology (such as strength of available materials) is progressing rapidly, and the market for space transport has expanded beyond the mostly government customer of the past to include a substantial commercial element. Selecting the optimal transportation system design is a primarily a function of the mission model, and secondarily a function of the risk level you choose and available capital you have to work with..

**Mission Model**

In space transportation system design, the "Mission Model" refers to
the information on what you want to transport in terms of quantity, size,
mass, type of cargo, etc. and when you want to transport it. A mission
model is developed from a project goal to define specific operating characteristics
that the transportation system must meet. For example, the Apollo
Program had a goal of landing a man on

the Moon before 1970. This is not sufficient information to design
a transportation system from. A mission model developed from this
goal would be something like:

__Cargo characteristics:__

Number of crew to the lunar surface: 2/mission

Maximum Stay time: 4 days/mission

Additional science equipment: 250 kg./flight

Lunar samples returned: 100 kg/flight

__Mission Schedule:__

First Flight: as early as possible but before Jan 1, 1970

Flight quantity: 10 to lunar surface (this was the original plan)

Flight rate: 4 flights/year

Some numbers will illustrate the problem. A good chemical rocket

has an exhaust velocity (the speed of the gases coming out the nozzle)

of 4500 m/s. The velocity to reach orbit is about 9000 m/s. The

basic equation of rocketry, the "rocket equation" tells you that

the ratio of rocket mass when full of fuel to rocket mass after

burning the fuel is:

m(i) / m(f) = exp ( dV / v(e) )

Where:

m(i) = intial mass

m(f) = final mass

dV = velocity change (9000 m/s in this case)

v(e) = exhaust velocity (4500 m/s in this case)

So in our example, dV/v(e) = 2, so m(i)/m(f) = exp(2) = 7.39.

Therefore 1/7.39, or 13.5% of the initial weight is left on reaching

orbit. In the past (before 1980s), the structure would be about 15%

of the takeoff mass, so there was a negative payload (i.e. you

couldn't get to orbit), even with a throw-away structure.

The rocket equation is generally valid for any type of reaction engine

with any velocity change.

In an attempt to increase the payload fraction, staging (dropping
part of

the rocket during the ascent) has been used. The vehicle is much lighter

as it burns off fuel. Less thrust, and hence fewer or smaller engines

are required in the later part of the launch. As propellant tanks are

emptied, they can be dropped off. A set of engines and tanks dropped
as

a unit is called a 'stage', and they are numbered in the order they
are

used and dropped (hence first stage, second stage, etc.). The drawback

to staging is that your vehicle must be re-assembled before the next

flight. This makes operating the vehicle more expensive.

To continue the example above, let us split the vehicle into two stages,

each of which provides half of the velocity to orbit. Using the rocket

equation, each stage has a ratio of initial to final mass, or mass
ratio,

of exp (1) = 2.72:1. Thus after the first stage burns it's fuel,

1/2.72 = 36.8% of the initial vehicle remains. The fuel for the first

stage represents 85% of the total first stage mass. The other 15%,

the structure and engines, is 11.1% of the total vehicle mass. So

the first stage in total is 74.4% of the total vehicle. The second

stage and payload is then 25.6% of the takeoff mass.

Similarly, the second stage has the same mass ratio, and so 36.8% of

it's mass is left after it burns it's fuel. Taking 15% for the structure,

we have 21.8% of the second stage+payload for the payload alone. Thus

the payload = 21.8% of 25.6% = 5.6% of the total vehicle mass. This

is a positive figure, unlike the single stage case, which is why all

rockets so far have used more than one stage.

The non-fuel mass of a stage can be grouped into engines, tanks,
and

'other'. Engines produce 40-100 times their weight in thrust. For

liftoff from the ground, you want about 1.3 times the vehicle weight

in thrust, so the engines are about 1.3-3% of the total weight. A

large tank, such as the Shuttle External Tank, can weigh 4% of the

fuel weight, but other tanks can range up to 10% of the fuel weight.

'Other' inlcudes plumbing, parachutes (if you want to use it again)

guidance systems, and such non-propulsion parts. It can range from

1% up to 10% of the total weight.

Older materials required 15% of the total weight for one-use

structures. Modern materials require about 10% of the total weight

for re-useable structures. Structures tend to get heavier at the

rate of 10% for each factor of 10 in life. So a 100-use structure

will be about 20% heavier than a one-use structure.

The circular orbit velocity, v(circ), for any body can be found
from:

v(circ) = sqrt ( GM/r )

Where:

G = Gravitational constant

M = Mass of body orbited

r = radius to center of body orbited

G is a univeral constant, and the mass of the Earth is essentially

constant (neglecting falling meteors and things we launch away from

Earth), so often the product G*M = K = 3.986 x 10^14 m^3/s^2 is used.

Escape velocity = sqrt ( 2GM/r ), or sqrt(2) = 1.414 times circular

orbit velocity.

Circular orbit velocity at the earth's surface is 7910 meter/sec.
At the

equator, the Earth rotates eastward at 465 meters/sec, so in theory
a

transportation system has to provide the difference, or 7445 meters/sec.

The Earth's atmosphere causes losses that add to the theoretical velocity

increment for many space transportation methods.

In the case of chemical rockets, they normally fly straight up intially,

so as to spend the least amount of time incurring aerodynamic drag.
The

vertical velocity thus achieved does not contribute to the circular
orbit

velocity (since they are perpendicular), so an optimized ascent trajectory

rather quickly pitches down from vertical towards the horizontal. Just

enough climb is used to clear the atmosphere and minimize aerodynamic
drag.

The rocket consumes fuel to climb vertically and to overcome drag, so
it

would achieve a higher final velocity in a drag and gravity free environment.

The velocity it would achieve under these conditions is called the
'ideal

velocity'. It is this value that the propulsion system is designed
to meet.

The 'real velocity' is what the rocket actually has left after the
drag and

gravity effects. These are called drag losses and gee losses respectively.

A real rocket has to provide about 9000 meters/sec to reach orbit,
so the

losses are about 1500 meters/sec, or a 20% penalty.

There is no law that says you have to use the same method of propulsion

all the way from the ground to orbit. In fact, it makes sense to use

different methods if one does better in the atmosphere and another
does

better in the later, vacuum part of the ascent.

In past rockets, this has been done by using different type of fuel

for different stages in a rocket. In the early part of the flight,
air

drag is important, so a dense fuel is preferred. A dense fuel means

smaller fuel tanks, and hence less area to create drag. Thus the

Saturn V used liquid oxygen/kerosine and the Shuttle uses solid rockets

for the first stage, both being dense fuels. Both use liquid oxygen/

liquid hydrogen for the second stage. This has the highest performance

in use for a chemical rocket fuel.

The Pegasus rocket uses an aircraft to get above the bulk of the

atmosphere. A sub-sonic jet engine has about ten times the performance

of a chemical rocket, mostly because it does not have to carry oxygen

to burn.

Many, many propulsion combinations are possible in getting to Earth

orbit and beyond. A large part of space propulsion design is choosing

which methods to use and when to switch from one to another.