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
version date comments
Various in house drafts
0.71 3 Jun 94 Translated from Word to ASCII and posted to
0.72 8 Jun 94 Cleaned up text, added ideas from Landis
0.77 Aug 99 Conversion to HTML files
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..
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:
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
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) )
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 )
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
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
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
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.