erichsen, peter spacecraft propulsion, a brief · pdf file · 2014-01-28the book...
TRANSCRIPT
Project coordination: Torsten H. Fransson Computerized Educational Platform Heat and Power Technology lecture series, Volume 13, 2
nd Edition
ERICHSEN, Peter
Spacecraft Propulsion, a Brief Introduction
Copyright © 2011 by Peter Erichsen
- i -
PREFACE FROM COMPEDUHPT
The “Computerized Education in Heat and Power Technology”
(=CompEduHPT) platform is designed as a fully electronic learning
and teaching platform for the field of Heat and Power Technology.
The project is a joint collaboration between persons involved in any
aspects of heat and power plant designs in a broad perspective,
including also any other kind of energy conversion, around the
world. The cluster is open to anyone who either contributes directly
with learning and/or teaching material or who is willing to sponsor
the development of the material in any other way.
Although all the “electronic books” existing in the project are
available inside the “CompEduHPT”-platform there has been a wish
from users as well as contributors that some of the books should also
be available in printed form, at a very reasonable price. This
“Lecture Series” is the outcome of this wish.
It is obvious that the CompEduHPT-platform, and its accompanying
Lecture Series, would not have appeared without the significant
interest in the project from all the CompEduHPT Cluster partners
worldwide. As initiator of the project I express my sincere thanks to
all my colleagues who have been willing to share their hard-earned
experience and learning/teaching material for the benefit of
“learners” around the world. I am also very grateful to all the
students (undergraduate as well as graduate) who have helped us to
develop the CompEduHPT-material to what it has become and where
it is heading. Needless to say that although the project would have
started without these persons, it would never have reached the
present state without their hard work.
Furthermore, the ideas and enthusiasm from these persons have
indicated that the vision of a fully interactive learning material
- ii -
corresponds to a future demand in the perspective of the life-long
learning.
I am also very grateful to the different organizations and companies
who have sponsored the work in various ways and at different times.
I hope that some of the results may be useful also to them.
The book “Spacecraft Propulsion Systems” has been written by Peter
Erichsen, formerly at the Swedish Space Corporation (SSC),
Sweden. It is based partly on SSC internal technical notes and on
course material that Peter Erichsen has been teaching at several
places over a number of years. It is with great pleasure that we
include this so far unpublished material in the CompEduHPT Lecture
Series.
Peter Erichsen has graciously agreed to share this material with the
CompEduHPT-platform on courtesy by SSC.
Torsten Fransson
Initiator of CompEduHPT-platform
PREFACE FOR 2nd EDITION
As its first issue, the new edition summarises propulsion
fundamentals as well as key features and performances of existing
and planned (near future) spacecraft propulsion systems. However,
with the introduction of the “System-specific Impulse”, Issp, as a
supplement to the rocket propulsion theory, this edition details also
the analysis of propulsion performances on spacecraft system level.
Peter Erichsen
- iii -
Table of Contents
1 INTRODUCTION ........................................................................................ 1
2 NEED FOR PROPULSION ......................................................................... 2
3 PROPULSION FUNDAMENTALS ............................................................ 4
3.1 BASIC PROPULSION EQUATIONS ................................................................. 4 3.2 PROPULSION PERFORMANCE ....................................................................... 7
3.2.1 Thruster Performance Factor ........................................................... 7 3.2.2 System Performance Factor ............................................................. 8 3.2.3 Evaluation of Mass of Propulsion Systems ..................................... 10
4 SURVEY OF SPACECRAFT PROPULSION SYSTEMS ................... 11
4.1 SPACECRAFT PROPULSION SYSTEM OPTIONS ........................................... 11 4.2 CHEMICAL PROPULSION ........................................................................... 17
4.2.1 Cold Gas ......................................................................................... 19 4.2.2 Hot Gas (survey) ............................................................................ 25 4.2.3 Monopropellant .............................................................................. 27 4.2.4 Bipropellant .................................................................................... 33 4.2.5 General System Design Considerations ......................................... 39 4.2.6 Solid Propellant .............................................................................. 40
4.3 ELECTRIC PROPULSION ............................................................................. 43 4.3.1 Propulsion Concepts ...................................................................... 43 4.3.2 Propulsion System Design and Performance ................................. 47
5 PROPULSION SYSTEMS SELECTION CRITERIA ............................ 57
6 OUTLINE OF POTENTIAL FUTURE SPACE PROPULSION ........... 64
6.1 POTENTIAL IMPROVEMENT OF CHEMICAL PROPULSION............................ 65 6.2 POTENTIAL IMPROVEMENT OF ELECTRIC PROPULSION ............................. 67 6.3 NEW APPROACHES.................................................................................... 69
7 GROUND TESTING OF PROPULSION SYSTEMS ............................. 72
8 MISSION SURVEILLANCE OF PROPULSION SYSTEMS ............... 73
9 LITERATURE/REFERENCES ................................................................ 74
-iv -
Standard Notations
Standard notation used throughout this booklet is given below.
C tank filling ratio (Vp/VT)
E energy [Ws]
F force [N]
g acceleration of gravity, standard, 9.81 [m/s2]
I impulse [Ns]
K tank performance factor (Pop/mT VT) [m2/s
2]
M molecular mass [kg/kmol]
m mass [kg]
m propellant mass flow rate [kg/s]
P power [W]
p pressure [N/m2]
R gas constant 8.314 [kJ/°K/kmol]
S/C spacecraft
T temperature [°K]
x non-impulse dependent system mass (%)
v velocity [m/s]
z gas compressibility factor
v velocity-increment m/s
specific power [W/kg]
overall power conversion efficiency (Pjet/P)
κ specific heat ratio
specific mass of propellant [kg/m3]
thrust time sec]
-v -
Subscripts
c motor chamber
case motor case
e exhaust (effective)
e-opt exhaust (optimal)
El electric (system)
H/W hardware
f final
jet thruster nozzle exhaust
0 initial
op operating
P propellant
PS propulsion system
PSS propellant storage system (tank with propellant)
S/C spacecraft
sp specific
ssp system-specific
T tank
tot total
- 1 -
1 INTRODUCTION
This booklet summarises key features and performance
characteristics of existing and planned (near future) rocket
propulsion for use on spacecraft such as satellites, space probes, etc.
In the frame of “Lecture Notes”, this booklet, presents a summary of
the “Rocket Propulsion Course” contained in the CompEduHTP-
platform with focus on spacecraft propulsion systems.
For a better understanding of spacecraft propulsion, the physical
background of propulsion is discussed and basic propulsion
mathematical equations are presented.
The aim of this presentation is to bring about the basics of space
propulsion on system level including propulsion system performance
evaluation. As a supplement to rocket propulsion theory, the
‘System-specific Impulse’ Issp is introduced. The Issp allows a more
accurate determination of the propulsive performance than the
commonly used ‘Specific-Impulse’ Isp which is only related to
propellant and thrust engine performances. The Issp has the advantage
in defining those parameters, which have a most significant impact
on propulsion system impulse performance. This allows
understanding the significance of the various system performance
parameters, which means also a better understanding of system
design concepts with related performance in general. Related exercises are noted under ‘Spacecraft Propulsion’ (S1B8C4) in the
CompEduHTP-platform: www.energy.kth.se/compedu.
An overview of basic common propulsion system designs is
presented together with tables and graphs which should allow the
valuation and facilitate a preliminary selection of propulsion systems
(chemical, electric) for spacecraft flight missions of given impulse
and velocity-increment requirement.
The literature noted in Chapter 9 is recommended for further reading
about spacecraft propulsion technology and its application.
- 2 -
2 NEED FOR PROPULSION
Propulsion is needed:
- to place payloads into orbit: launch propulsion;
- to send payloads to the moon or to the planets: space
propulsion;
- to position, adjust and maintain orbits of spacecrafts by orbit
control: auxiliary propulsion;
- to orient spacecraft by attitude control: auxiliary propulsion
also called reaction-control systems.
There are the following types of reaction-control systems:
- reaction jets (propulsion): which produce a control force by
the expenditure of mass;
- solar sails, magnetic torquers (magnetic coils): which produce
a control force by interaction with the environmental field;
- momentum-transfer devices (reaction-, flywheels): which
produce no net control force, but simply transfer angular
momentum to or from the spacecraft.
In this booklet, only propulsion systems will be dealt with which are
based on jet propulsion devices that produce thrust by ejecting stored
matter, called the propellant.
- 3 -
The main features of jet propulsion are:
a) LAUNCH PROPULSION for launching rockets with the
following characteristics:
- high velocity increment capability (7 - 11 km/s)
- very high thrust levels (ratio thrust/launch vehicle weight: 1.3)
- low fraction of launch vehicle take-off mass for payload (1 - 5%)
- powerful chemical rockets
b) SPACECRAFT PROPULSION is characterised in general by its
complete integration within the spacecraft. Its function is to
provide forces and torques to:
- transfer the spacecraft: orbit transfer incl. interplanetary travel
- position the spacecraft: orbit control
- orient the spacecraft: attitude control
While jet propulsion systems for launching rockets are also called
primary propulsion systems, spacecraft, e.g. satellites, are operated
by secondary propulsion systems.
In order to fulfil attitude and orbit operational requirements,
spacecraft propulsion systems are characterised in particular by:
- low thrust levels (1 mN to 500N) with low acceleration levels,
- continuous operation mode for orbit control,
- pulsed operation mode for attitude control,
- predictable, accurate and repeatable performance (impulse bits),
- reliable, leak-free long time operation (storable propellants),
- minimum and predictable exhaust plume impingement effect.
- 4 -
3 PROPULSION FUNDAMENTALS
3.1 Basic Propulsion Equations
The essence of space propulsion is to modify the velocity vector of a
spacecraft either in magnitude or in direction so as to modify the
orbit or attitude. However, an isolated body, like a spacecraft, can
modify its momentum only if external forces act on it, since all
internal forces cancel each other in action-reaction pairs. This is
expressed by Newton's law of motion:
dt
dmv
dt
vdm
dt
vmdF
)(, (1)
Unfortunately, there are no such external forces in space (except very
weak perturbation forces which have to be compensated by the
spacecraft onboard propulsion system) and Eq. (1) becomes:
0dt
dmv
dt
vdm
(2)
Therefore, the only and obvious way out is, that the spacecraft must
be split up such, that a part of the spacecraft can modify its velocity
through the effect of action-reaction forces.
In fact, this is realised by the ejection of mass in form of propellant
from the spacecraft. If we assume a spacecraft with a mass m,
ejecting propellant with a rate of dm/dt at constant velocity evv
at
nozzle outlet, Eq. (2) can be written with the action and reaction
forces in balance:
dt
dmv
dt
vdm e
(3)
- 5 -
This expresses that the spacecraft experiences acceleration in the
opposite direction to ev, or that the external force acting on the
spacecraft is by definition the force of thrust. That is, constant
exhaust propellant velocity evv
at nozzle outlet ( evv
is the
relative velocity between spacecraft and exhaust propellant) gives the
basic equation for force of thrust:
Fdm
dtv mve e
N, (4)
with dm
dtm
s
kg for the propellant mass flow rate. (5)
Strictly speaking, ev and F are vectors. They are here taken to be
collinear, so no vector notation is needed.
Eq. (3) can be integrated to get the accumulated velocity increment
v of a spacecraft:
dv vdm
mv
v v
e
m
m
o
f
0
. (6)
Integrated:
v vm
me
f
o
ln
s
m, (7)
where m0 is the initial mass of the spacecraft at the beginning and mf
is the final mass of the spacecraft at the end of its mission.
- 6 -
This formula can be also written in the form of the basic 'Rocket
Equation':
m
me
f
v
ve
0
(Tsiolkovsky-Equation). (8)
The propellant quantity required for a spacecraft velocity change v
is with mf = m0 - mP:
)1(0ev
v
P emm
kg . (9)
Some other useful definitions
The total impulse delivered by a certain quantity of propellant is
calculated by:
Pe
m
etot mvdmvFdtIP
00
Ns . (10)
The kinetic energy of ejected matter is:
2
2
1ePjet vmE Ws . (11)
And the power of the jet is calculated by:
PdE
dtmv F
vjet
jet
e
e 1
2 2
2
W . (12)
Therefore, the power input for an electric thruster will be:
PP
mv
Fvjet e e
2
2 2 W , (13)
where is the power conversion efficiency.
For further reading about propulsion fundamentals see [1] and [2].
- 7 -
3.2 Propulsion Performance
3.2.1 Thruster Performance Factor
The most useful parameter for determining thrust engine (or thruster)
performance is specific impulse:
IF
msp
kg
Ns, (14)
This is defined as the impulse delivered per unit mass of propellant
and which can be easily obtained by test, i.e. by measuring of the
thrust F and propellant mass flow rate
m with help of a thrust stand
in a vacuum chamber; see Fig. 1. THRUST AND SPEC. IMPULSE MEASUREMENT
F - signal
Accuracy of Measurements (typical)
Pressure - P (bar) 0.2 %
Temperature - T (o C) 2 %
Propellant Mass Flow -
m (g/s) 0.3%
Thrust - F (N) 0.2%
Spec. Impulse - Isp (Ns/kg) 0.5%
Thruster
Thrust Stand
F
Vacuum Chamber
m - signal T - signal
P- signal
m
Fig. 1: Measurement of Thrust F and Propellant Mass Flow Rate
m
- 8 -
An effective exhaust velocity of the jet is introduced, which is
determined by test:
vF
me
s
m. (15)
From its definition as the thrust per unit rate of mass flow of
propellant, it follows that ve is numerical the same as the Isp as
defined above with SI units of m/s. Note: ve hereafter is always the
effective exhaust velocity, although called simply ‘exhaust velocity’,
if not stated otherwise. Further, in all related calculations with ve, the
effective velocity has to be applied.
Propulsive performance is commonly associated with the specific
Impulse Isp, (ve). According to the ‘Rocket Equation’ (8), a high value
of Isp will result in a mission final high spacecraft mass, which means
high payload mass, because of lower propellant mass consumption
during the spacecraft mission.
Specific impulses are sometimes quoted in units of seconds,
corresponding to a modification of the above definition to that of the
impulse delivered per unit weight of propellant. Such values in
seconds then follow from those in Ns/kg by division with the
gravitational acceleration standard, g (= 9.8 m/s2).
3.2.2 System Performance Factor
With regard to the evaluation of propulsion performance on system
level, propellant storage systems, and especially for electric
propulsion, electric power supply and power processing systems (see
Chapter 4.3.2) may form a major ‘dead’ dry mass of the overall
propulsion system mass. Therefore, the choice and sizing of
propulsion systems is not always clear on the basis of Isp alone.
- 9 -
In general, it can be assumed that, especially for missions with high
total propulsion impulse (e.g. geostationary and interplanetary
missions), the mass of the corresponding auxiliary propulsion system
may represent an important fraction of the overall mass of the
spacecraft. Attempts to minimise the mass of propulsion systems
have therefore to concentrate on parameters, which characterise the
system’s propulsive performance capabilities. Hence, a system
reference number has to be defined, describing those design
parameters which influence propulsion system mass in relation to
delivered impulse.
To describe the performance of the entire spacecraft propulsion
system, a reference number is introduced, which defines the total
impulse Itot, delivered by the entire propulsion system mass mPS:
II
mssp
tot
PS
kg
Ns. (16)
Because of the resulting dimension, - delivered impulse per kilogram
of system mass mPS, this number is called System-specific Impulse,
[3].
The Issp can be directly derived from actual spacecraft propulsion
systems by determining the total impulse delivered by the system
contained propellant (see Eq. (10)), divided by the mass of the
propulsion system (including mass of contained propellant).
On the other hand, the Issp can be derived analytically. The Issp varies
with the kind and design of propulsion systems. In Chapter 4 below
the most common spacecraft propulsion systems are presented with
derived Issp mathematical formulas and with relevant propulsion data.
- 10 -
3.2.3 Evaluation of Mass of Propulsion Systems
The mass of propulsion systems can be derived from the propulsion
system mass fraction. The dependence of the propulsion system mass
fraction on mission velocity increment v is derived from the
‘Rocket Equation’ in combination with the definition of the system-
specific impulse Issp.
The first Eq. (17) below is obtained from Eqs. (9) and (10). The
second Eq. (18) is just the definition of Issp, and the final expression
Eq. (19), follows from the first two:
)1(/ev
v
CSePetot emvmvI
(17)
PSssptot
PS
tot
ssp mIIm
II (18)
)1()1(/
ee v
v
ssp
spv
v
ssp
e
CS
PS eI
Ie
I
v
m
m
(19) (19)
where mS/C m0 is the (initial) mass of the spacecraft and with the
understanding of:
kg
NsI
s
mv spe is numerical equal. (20)
With the help Eq. (19), for given values of ve and Issp, the propulsion
system mass fraction can be plotted as a function of velocity
increment (v), as presented in Chapter 5 and realised by the
computerised ‘Issp-Program’, see [4]. By this, the Issp and the mass
fraction of propulsion systems can be evaluated for given mission
impulse and v requirements.
- 11 -
4 SURVEY OF SPACECRAFT PROPULSION SYSTEMS
4.1 Spacecraft Propulsion System Options
Spacecraft propulsion systems can be classified according to the type
of energy source. Both, space propulsion and auxiliary propulsion are
performed by the following two main on-board spacecraft propulsion
system types:
A) Propulsion Systems with self-contained energy in propellants, comprising cold gas and hot gas systems. The energy to
produce thrust is stored in the propellant, which is released
mainly by chemical reactions (this is why these systems are
mostly referred to as chemical propulsion systems) and the
propellant is then accelerated to a high velocity by expanding it
in form of gas through a nozzle. These systems contain:
- Storage and feed system that stores (tank) and feeds the
propellant to the thrusters to generate thrust.
- Valves, piping which connects the propellant storage system
with the thruster.
- Electric control unit to operate electrically the valves and
thrusters.
Thrust Exhaust
Propellant Storage and
Feed System
Thrusters, Valves,
Piping, etc.
Electrical
Control Unit
Figure 2: Schematic of Chemical Propulsion Systems
- 12 -
With regard to the system-specific impulse, Issp = Itot/mPS (see
Eq. (16) above), its practical application, especially for system
performance analysis, requires a very clear definition of what is
included in the total mass of propulsion system mPS.
Therefore, the Issp can be further detailed according to Fig. 2, and
with Eq. (16) the Issp for chemical propulsion systems can be
written:
PSSWH
tot
sspmm
II
/
, (21)
with mH/W, the propulsion hardware mass, such as thrusters,
valves, piping, etc., which is independent of propulsion impulse,
and mPSS, the mass of propellant storage system (propellant +
tank), which is proportional to propulsion impulse; see coloured
box of Fig. 2.
The following types of propulsion systems are part of systems
with self-contained energy in propellants:
- Cold gas systems, comprising inert gases (e.g. nitrogen: N2)
and high vapour pressure hydrocarbons (e.g. ammonia:NH3
and propane: C3H8).
- Monopropellant hydrazine systems (N2H4).
- Storable bipropellant systems (e.g. nitrogen tetroxide (NTO:
N2O4) oxidiser with anhydrous hydrazine (N2H4) or
monomethyl hydrazine (MMH: CH3N2H3) fuels).
- Solid propellant motors (composite propellants: e.g.
aluminium powder with hydroxyl-terminated polybutadiene
(HTPB) binder and an oxidiser like ammonium perchlorate).
More details about chemical propulsion are presented in Chapters 4.2
below.
- 13 -
B) Propulsion Systems with externally supplied energy to
propellant, comprising e.g. electric propulsion. The energy to
produce thrust is not stored in the propellant but has to be
supplied from outside by an extra power source, e.g. nuclear,
solar radiation receivers (solar cells) or batteries. These systems
contain:
- Storage and feed system that stores and feeds the propellant
to the thrusters to generate thrust.
- Valves, piping which connects the propellant storage system
with the thruster.
- Electric control unit to operate electrically the valves and
thrusters.
- Electric power generator and power processing system.
Electrical Power
Generator
Power Processing
System
Control Unit, Harness,
Piping, etc.
Propellant Storage and
Feed System
Electrical Thruster
Assembly
Plasma/
Ion Jet
Figure 3: Schematic of Electrical Propulsion Systems
- 14 -
According to Fig. 3 and with Eq. (16) the Issp for electric
propulsion systems can be written:
ElPSSWH
tot
sspmmm
II
/
(22)
with mEl, to be added to the system mass with regard to chemical
propulsion. The mEl comprise the mass the electrical power
generator, the power processing system and the electrical
thrusters assembly which are proportional to the power to be
handled by electric propulsion systems; see coloured box of
Fig. 3.
The following types of propulsion systems are part of systems with
externally supplied energy to propellant, i.e. electric propulsion:
- Electrothermal systems (resistojet and arc-jets):
Here thrust is produced by expansion of hot gas (which is heated
by electric current) in a nozzle.
- Electromagnetic systems (magnetoplasmadynamic: MPD).
- Electrostatic systems (ion engines: Kaufman, radio-frequency,
field emission, stationary plasma):
Here thrust is produced by acceleration of charged particles in
electric or magnetic fields to high expulsion velocities.
More details about electric propulsion are presented in Chapters 4.3
below.
Eqs. (21) and (22) for Issp of the various propulsion systems have in
common the same numerator, representing the total impulse
- 15 -
delivered by the propellant contained in the propellant tank, which is
with Eq. (17):
ePtot vmI (23)
while the denominator in Eqs. (21) and (22), with regard to the
impulse related system mass (mPSS, mEl), varies with the kind and
design of propulsion systems. In this respect, a concise description
of common spacecraft propulsion systems is presented below.
Details of derived mathematical formulas of Issp for chemical
propulsion systems are presented in Chapter 4.2 and those for
electrical in Chapter 4.3 below.
An overview of actual propulsion system options according to their
source of energy is shown in Fig. 4.
Figure 4: Classification of Spacecraft Propulsion Systems
Classification of Propulsion systems
PROPULSION SYSTEMS
CHEMICAL
COLD GAS HOT GASELECTROTHERMAL
(Resistojet; Arcjet)
ELECTROMAGNETIC
( MPD-Thruster)
ELECTROSTATIC
(RIT; Field emission) VAPORISING LIQUID
COMPRESSED GAS
(Nitrogen)
(Propane)
SOLID PROPELLANT
MONO-PROPELLANT
(Hydrazine)
BI-PROPELLANT
(MMH/N2O)
ELECTRICAL
- 16 -
Historically, chemical propulsion, comprising cold gas and hot gas
systems, was the first one available for space propulsion which is
now followed by the development of electric propulsion systems.
Presentations of spacecraft propulsion systems in Chapter 4 below
will be concentrated on today’s most commonly used system
designs, which include traditional chemical propulsion with evolving
environmental benign, so-called ‘green propellants’ as well as
electric propulsion still under development.
Finally, Chapter 6 will present an outline of the potential future
evolution of spacecraft propulsion.
- 17 -
4.2 Chemical Propulsion
Chemical propulsion is based on the principle of converting chemical
energy (or pressure) contained in the propellant to kinetic energy of
thrust engine exhaust gases.
Currently available chemical propulsion systems can be categorised
as either hot gas, or as a border-line case, cold gas.
Propulsion operating with cold gas, represents the simplest form of a
propulsion system, It comprise compressed (inert) gas which is
stored at high pressures in a tank, and vaporising liquids (high
vapour pressure hydrocarbons), which are pressurised by their own
equilibrium vapour pressure. Expelling these gases through a nozzle
creates a thrust force.
Propulsion systems operating with hot gas comprise systems
containing liquid and solid propellants. The energy from an
exothermal combustion reaction of the propellant chemicals in a
thruster results in high temperature reaction product gases, which are
expelled through a nozzle. The maximum exhaust velocity will be
achieved when all enthalpy contained in the gas at the inlet of the
nozzle is transferred into kinetic energy by its expansion in the
nozzle. This is described by the equation of ‘Saint-Venant’ for an
ideal nozzle with a complete expansion of the gas at the outlet of the
nozzle:
M
RTve
)1(
2max
, => in general: ve
M
T , (24)
with the assumption that besides R (gas constant) also κ (specific
heat ratio) is constant. Therefore, for high values of ve, high gas
temperatures T and low molecular mass M are required.
However, it has to be noted that nozzles are not of infinite length so
that gases are not expanded down to absolute vacuum and therefore
- 18 -
gases leave the outlet of the nozzle with residual enthalpy. In
addition, exhaust velocities ve will be limited by nozzles wall friction
loss, jet divergences, condensation of gas if temperatures become
low enough.
Typical values of ve/ve max for chemical propulsion systems are about:
0.85 ÷ 0,95 for cold gas systems with no thermal losses and
very high area ratio nozzles thus higher nozzle expansion
ratios.
0.6 ÷ 0.8 for hot gas systems because of heat losses.
Finally, exhaust velocities ve will be limited by the available energy
release per unit of mass of propellant which is according to Eq. (11):
P
em
Ev
2 (25)
One of the most energetic chemical reactions release energies such as
for O2 + H2 is about 13.4106 Joules/kg and ve is then 5200 m/s
theoretically, while real values are being around ve = 4200 ÷ 4500
m/s. Therefore, for chemical propulsion, maximum jet exhaust
velocities are limited to <5000 m/s.
Terminology:
Rockets using solid propellants are called motors.
Rockets using liquid propellants are called engines.
The term thruster is used for small thrust application, e.g.
spacecraft auxiliary propulsion systems.
- 19 -
4.2.1 Cold Gas Cold gas systems operate with compressed inert gas (e.g. nitrogen:
N2) or high vapour pressure hydrocarbons (e.g. ammonia, NH3); see
Table 3 below.
Cold gas systems are shown schematically in Fig.5. The typical
system consists of a propellant tank, fill valve, filter, pressure
regulator, line pressure transducers, control valves, and nozzles. The
pressure regulator provides propellant at constant pressure as the tank
pressure drops. A relief valve is incorporated downstream of the
pressure regulator to prevent system rupture in the case of a regulator
failure. With regard to compressed gas systems, the cold gas is stored
at high pressures (200 - 300 bar) in a tank.
Figure 5: Basic Flow Scheme of Cold Gas Propulsion Systems
- 20 -
The vaporising liquid system is characterised by a liquid propellant
pressurised by its own equilibrium vapour pressure and the expulsion
of this vapour through a nozzle. In order to provide completely
vaporised gas, a vaporiser is included in liquid cold gas systems.
A typical cold gas thruster configuration is shown schematically in
Fig. 6 below.
Figure 6: Cold Gas Thruster Configuration
A cold gas thruster consists of a solenoid valve with mounted nozzle.
The thruster is operated by opening the solenoid valve with help of
an electric current. Typical thrust range is 0.02 to 10N for spacecraft
attitude and orbit control.
System-specific Impulse, Issp
With regard to the derivation of the Issp for cold gas systems, the
denominator of Eq. (21) has to be further evaluated.
Starting with compressed cold gas systems, usually the cold gas
used is stored at high pressures in a tank. Therefore, for calculating
the gas mass content in the tank, the gas law applies as follows:
TM
RzmVp PTop (26)
- 21 -
For calculating the tank mass, the so-called Tank- Performance
Factor usually defined as:
T
Top
m
VpK is to be used. (27)
Table 1 below presents typical values of tank K-factors for different
built tank designs and different tank materials for compressed gas,
[5]. According to Eq. (27), the higher the K-factor, the lower will be
the mass of the tank. Consequently, the tank material shall be a high
tensile strength material, such as Titanium Alloy or even better, a
fibre composite material, like Kevlar, see Table 1 below. With regard
to the tank safety factor, see Chapter 4.2.5.
Table 1: Ranges of typical Propellant Tank Performance Factors, K,
for High Pressure Tanks
Type of Tank Average K*
(104 m2/s2)
Range
(+/- 1sigma)
(104 m2/s2)
Remarks
*Tank Safety
Factor: S=2
Titanium Alloy: Ti 6Al4V 6.43 5.87-6.99 High Pressure
Composite Over-wrapped
Pressure Vessel
12.20 8.29-16.11 Tanks
With Eq. (26) and (27) the combined mass of the tank and propellant
is:
KM
zRTmmmm PTPPSS 1 (28)
- 22 -
With regard to the non-impulse dependent system mass, mHW, such as
thrusters, valves, piping etc., if properly known, mHW can be included
as a mass fraction x (%) of the impulse dependent system mass mPSS.
Table 2 presents values of x based on built spacecraft, which are
noted according to the class of spacecraft. Note, the larger the
spacecraft, in general the larger will be also the impulse dependent
part of the propulsion system because of higher impulse demands
due to higher spacecraft mass. Therefore, the mass fraction x tends to
get lower with respect to increasing tank and propellant mass.
Table 2: Non-Impulse System Mass Factor, x (%), [6]
Class of satellites Nano Micro Mini
(Small satellites)
Medium sized
Macro
(Large satellites)
Mass of satellites [kg] (indication)
1 – 10 10 – 100 100- 500 500 – 1000 > 1000
Factor ‘x’ [%] 21 24 5.6 6.3 4.5
(Average x values of examples for different satellites classes are only indicative)
Hence, Eq. (28) can be expanded to include mHW:
xKM
zRTmmmmmm PHWTPHWPSS
11 (29)
Consequently, with Eqs. (21), (23) and (29), the system-specific
impulse for COMPRESSED GASES becomes:
xKM
zRT
v
mmm
II e
HWTP
totssp
11
(30)
- 23 -
Further, for vaporising liquids, the mass of liquids in a tank is:
PP Vm (31)
In order to allow certain ullage, the volume of propellant is a certain
fraction (0.5 to 0.9) of the available tank volume. Hence, with VP =
CVT and Eqs. (27) and (31) for the combined mass of tank and
propellant we get:
KC
pmmmm
op
PPTPSS
1 (32)
Again, with regard to the non-impulse dependent system mass, mHW,
as for compressed gas systems, it can be included as a mass fraction
x (%) of the impulse dependent system mass mPSS. Hence, Eq. (32)
can be expanded to include mHW:
xKC
pmmmmmm
op
PHWPTHWPSS
11
(33)
With Eqs. (21), (23) and (33) the system-specific impulse for
VAPORISING LIQUIDS becomes:
xKC
p
v
mmm
II
op
e
HWTP
tot
ssp
11
(34)
From Eqs. (30) and (34) it is obvious, that the thruster exhaust
velocity, ve ≡ Isp, the non-impulse dependent system mass x as well
as the type of propellant and the tank performance factor K
- 24 -
influences the Issp. However, high values of Issp are mainly dictated
by high values of ve and low values of x, while all other parameters
are of secondary importance.
With regard to vaporising liquids, while no great improvement over
inert gas thrusters exhaust velocity ve can be obtained, considerable
savings in propellant storage mass result from the propellant’s high
density and low pressure. This is illustrated by values of Isp and Issp
as in indicated by examples for actual cold gas propulsion systems,
presented in Table 3 below. Note, that calculated values of Issp show
a good agreement with actual values of Issp. This confirms that the
Issp-analytical tool describes very well those design parameters which
characterise the system’s propulsive performances.
Table 3: Actual Cold Gas Propulsion Systems Performances (Listed data are examples and therefore only indicative)
PROPELLANT
THRUSTER
SPEC.-
IMPULSE
Isp (mission
average)
(Ns/kg)
TOTAL
IMPULSE
Itot
(Ns)
PROP.
SYSTEM
MASS
mPS
(kg)
SYSTEM
SPEC.-
IMPULSE
Issp
(Ns/kg)
REMARKS
Actual Propulsion Systems
Issp values derived from
Ref. 7 if not noted
otherwise
Nitrogen (N2) 706 845 4.4 193 Vela III
Mol.Mass (M):
28 kg/kmol 706 6780 24 283 COS-B; 8
z = 1.13, at tank
pressure 250 bar 706 - - 273 Calculated: tank material:
Ti 6Al 4V; K= 6.8·104
m2/s
2
x = 5.6% (small S/C)
Argon (A) 510 3900 16.8 232 OGO A,B,C
Mol.Mass (M):
39.9 kg/kmol 510 5500 24.3 226 TD-1A 9
z = 1.02, at tank
pressure 250 bar 510 - - 248 Calculated: tank material:
Ti 6Al 4V; K= 6.8·104
m2/s
2
x = 5.6% (small S/C)
Ammonia (NH3) 800 4450 6.8 654 NRL Explorer 30 (1965)
=0.62 (kg/m3)·10
3
Max. op. pressure
at 30C: 12 bar
800 - - 663
Calculated: Al tank with
vaporizer: K=1.5·104
m2/s
2
C=0.9; x = 5.6% (small S/C)
- 25 -
Conclusion
Although of moderate impulse capability, cold gas systems, in
particular systems operating with compressed cold gas, as with
Nitrogen, N2, are still of interest in view of their simplicity, high
reliability and repeatability of impulse bit.
ADVANTAGES:
- simplicity and reliability;
- lowest cost propulsion system;
- very low thrust ( 10 N) and impulse bit ( 10-4
Ns)
capability;
- low plume contamination.
DISADVANTAGES:
- low Isp ( 950 Ns/kg) low Issp ( 650 Ns/kg) with
resulting high system mass.
4.2.2 Hot Gas (survey)
For increasing absolute levels of thrust and impulse requirements for
spacecraft propulsion (e.g. orbit transfer and orbit control), cold gas
systems are inadequate and more energetic propellants generating hot
gas for mass expulsion are required, see Eq. (24).
Hot gas systems are the most common type of propulsion systems for
space applications. They can be divided into three basic categories:
- liquid, comprising monopropellant hydrazine (N2H4) and
storable bipropellant (MMH/N2O4);
- solid, with composite propellants;
- hybrid (so far not used for spacecraft propulsion).
- 26 -
The terminology refers to the physical state of the stored propellants
as illustrated in Fig.7 below. Note, only systems containing liquid in
form of monopropellant hydrazine or bi-propellants and solid
propellants are used for spacecraft propulsion, while hybrid
propulsion systems are used for launch propulsion.
Figure 7: Schematic of Hot Gas Propulsion Systems
In contrast to compressed gas and vaporising liquids, liquid
propellants need to be pressurised in the tank to feed the thrusters
with propellant. Note that due to long spaceflight mission durations,
only pressure-fed systems are used because of their inherent
simplicity compared with pump-fed systems, which are used
commonly for launch propulsion. Therefore, hot gas propulsion
systems for spacecrafts in the gravity-free environment need
propellant tanks equipped with propellant management devices in
order to separate liquids from the pressurising gas; - details see
Chapters 4.2.3 and 4.2.4 below.
Oxidiser
Oxidiser (liquid)
Fuel
Fuel (solid)
Liquid
Hybrid
Solid
Solid Propellant
- 27 -
4.2.3 Monopropellant
Monopropellant systems use a single (Mono) propellant to produce
thrust. The most commonly used monopropellant is anhydrous
hydrazine (N2H4), as noted in Table 5 below. The hydrazine
propellant is decomposed in a thruster by a catalyst and the resulting
hot gas is expelled through a nozzle, thus generating thrust force on
the spacecraft. A typical monopropellant system, as shown
schematically in Fig.8, uses generally nitrogen gas to expel the
propellant from a diaphragm tank into the chamber catalyst beds of
the thrusters. The typical system contains fill and drain valves for the
pressurant gas and for the monopropellant hydrazine.
Figure 8: Basic Flow Scheme of Monopropellant Hydrazine
Propulsion Systems
T
P
Hydrazine System with Catalytic Thrusters
T
Hydrazine Gas Generator System
vvvvvvv
T T
X
Fill Valve (Nitrogen)
Temperature Sensor
Propellant Tank(diaphragm)
Fill Valve (Hydrazine)
Pressure Transducer
Filter
Thruster
P
P
Latch-
Valve
SolenoidValve
HydrazineDecomposit.Chamber
Gas-Store
Tank
Filter
N2H4
N2 N2
N2H4
T
P
Hydrazine System with Catalytic Thrusters
T
Hydrazine Gas Generator System
vvvvvvv
T T
X
Fill Valve (Nitrogen)
Temperature Sensor
Propellant Tank(diaphragm)
Fill Valve (Hydrazine)
Pressure Transducer
Filter
Thruster
P
P
Latch-
Valve
SolenoidValve
HydrazineDecomposit.Chamber
Gas-Store
Tank
Filter
N2H4
N2 N2
N2H4
- 28 -
In addition the system contains latch valves and line pressure
transducers. Filters are provided upstream of line valves to prevent
damage of the valve seats or clogging the injectors of thrusters by
entrained foreign material.
Since the pressurant gas is stored (at a pre-selected but relatively low
pressure, e.g. 22 bar) in the propellant tank, the propellant pressure
varies with propellant usage. A typical selection of the ullage volume
of 25% filled with pressurant gas (thus containing 75% propellant)
will result in a propellant feed pressure decay, and thus in a thrust
decay of 4:1. This mode of operation is also referred to as the blow-
down mode, in contrast to the pressure constant mode, which
requires the storage of a high-pressure gas in a tank external to the
propellant tank (see ‘Bipropellant systems’).
A typical monopropellant hydrazine thruster configuration is shown
schematically in Fig.9 below. Thrust is produced by decomposition
of hydrazine into hot gas in the presence of a catalyst such as iridium
metal supported by high-surface-area aluminum oxide granulates.
The catalyst causes the hydrazine to decompose into ammonia,
nitrogen gas and hydrogen gas at high temperature up to 1100 oC.
This results in a fairly high specific impulse of up to Isp= 2300 Ns/kg.
Typical thrust range is 0.5 to 22N for spacecraft attitude and orbit
control maneuvers.
Propellant Inlet with Filter
Solenoid Valve (not shown)
Heat Barrier
Injector Head
Injector
with Catalyst
Nozzle
Decomposition Chamber
Thermal Insolation
Figure 9: Monopropellant Hydrazine Thruster Configuration
- 29 -
System-specific Impulse, Issp
With regard to the derivation of the Issp for systems operating with
liquid propellants, the denominator of Eq. (21) has to be further
determined. In contrast to compressed gas and vaporising liquids,
liquid propellants need to be pressurised to feed the thrusters with
propellant. Therefore, the mass of the pressurising gas mpr and, if
necessary, an extra tank with mTpr for the pressurising gas has to be
taken into account.
No extra tank for the pressurising gas is needed for the blow-down
mode, which is the most widely used means of tank pressurisation
for monopropellant hydrazine. As already mentioned above, at the
beginning of a mission the volume of the propellant is a certain
fraction C (mostly 0.75 for a blow-down ratio of 4:1) of the internal
tank volume. Consequently, the volume of pressurising gas in the
propellant tank will be Vpr = (1-C)VT. Therefore, the mass of the gas
can be derived easily from the gas law and will be with Eqs. (26) and
(27):
zRT
MCKmm T
pr
)1( (35)
With Eq. (33) and (35) the combined mass of tank with propellant
incl. pressurising gas and non-impulse dependent system mass mHW
is given by:
HWprPTHWPSS mmmmmm =
= xzRT
MCK
KC
pm
op
P
1
)1(11
(36)
- 30 -
With (21), (23) and (36) we obtain the final expression for the Issp of
systems operating with stored liquid and with contained pressurising
gas in the propellant tank, representing the ”BLOW-DOWN MODE”:
HWprPT
totssp mmmm
II
=
xzRT
MCK
KC
p
v
op
e
1)1(
11
(37)
Eq. (37) shows, that both the type of propellant (represented by ve, ,
- a high ve and a high are desirable) and the propellant storage
conditions (propellant-storage pressure pop, tank-filling ratio C, type
of pressurising gas, tank performance factor K) influence the system-
spec. impulse. However, as already mentioned for cold gas, it has to
be noted from Eq. (37), that high values of Issp are mainly dictated by
maximum values of thrusters exhaust velocity ve (Isp) and low values
of impulse independent system mass x, while all other parameters
noted above will have only a secondary impact on values of Issp.
With regard to tank K-factors, propellant tanks, with liquid
propellants in contrast to cold gas systems, need propellant
management devices. In the case of monopropellant hydrazine, tanks
are equipped with positive expulsion devices, which are diaphragms,
and which mechanically separate the pressurizing gas from the liquid
propellant in the tank during the gravity-free condition of spaceflight
missions. Diaphragms are made typically of Buthyl-and Ethylene
Propylene rubber materials. For more aggressive bi-propellant
liquids (see Chapter 4.2.4. below), only surface tension devices made
of stainless steel screens can be used. They work by using surface
tension forces between the propellant liquid and the metal screen to
separate liquid from the pressurising gas.
- 31 -
Table 4: Ranges of typical Propellant Tank Performance Factors-K
for Liquid Propellant Tanks, [5]
Type of Tank Average K*
(104 m2/s2)
Range
(+/-1 sigma)
(104 m2/s2)
Remarks
* Tank Safety
Factor: S = 2
Diaphragm 2.25 1.53-2.97 Tank Material:
Surface tension 3.32 2.28-4.36 Titanium Alloy
No propellant propellant
management device
4.06 3.41-4.71 Ti 6Al4V
Because of the more energetic propellant hydrazine generating hot
gas for mass expulsion, higher values of ve, and Issp are achieved
when compared with cold gas thrusters; see Table 5 below.
Table 5: Performances of Actual Hydrazine Propulsion Systems
(Listed data are examples and therefore only indicative)
PROPELLANT
THRUSTER
SPEC.-
IMPULSE
Isp (mission
average)
(Ns/kg)
TOTAL
IMPULSE
Itot
(Ns)
PROPUL-
SION
SYSTEM
MASS
mPS
(kg)
SYSTEM
SPEC.-
IMPULSE
Issp
(Ns/kg)
REMARKS
Actual Propulsion
Systems
Issp values derived
from Ref. 10
Monopropellant 2163 2.64 105 142 1859 ECS
Hydrazine: 2134 6.40 105 375 1707 ERS-1
N2H4 2110 9.50 104 66 1440 EXOSAT 2110 6.41 104 38 1687 GEOS
ve 2300 m/s; 2163 1.49 105 80.1 1860 GIOTTO
= 1.0·103 kg/m
3 2163 7.25 104 48.8 1486 ULYSSES
2168 2.36 105 130 1815 MARECS
2060 8.24 104 53 1555 METEOSAT
2168 3.04 105 168 1810 TELECOM-1
N2 pressurant gas,
Max. op. pressure:
pop=22 bar
2150 − − 1781 Calculated: Issp for
diaphragm tank:
C=0.75,
K=2.3104 m2/s2,
x = 6% (medium S/C)
- 32 -
For comparison, a calculated Issp-value for typical propellant storage
and non-impulse system mass x, parameters are also presented in
Table 5, showing an overall good agreement with Issp-values of actual
spacecraft propulsion systems.
Conclusion
Monopropellant hydrazine for spacecraft attitude and orbit control is
one of the most widely used propellants. The primary reason for such
wide acceptance of monopropellant hydrazine propulsion systems
lies in their inherent simplicity (reliability) while still providing
adequate propulsive performance.
ADVANTAGES:
- simplicity and reliability (monopropellant);
- lowest cost propulsion system (other than cold gas);
- space storable for long periods (> 15 years demonstrated);
- low thrust capability ( 0.5 N);
- moderate thrust levels available ( 22 N).
DISADVANTAGES:
- moderate Isp ( 2300 Ns/kg) with moderate Issp
( 1900 Ns/kg) resulting in medium to high system mass;
- limited life of catalyst.
- 33 -
4.2.4 Bipropellant
Bipropellant systems are characterised by the combustion of two
(Bi) propellants, a fuel (e.g. monomethyl-hydrazine, CH3NHNH2)
and an oxidiser (e.g. nitrogen tetroxide, N2O4), to produce thrust.
A typical bipropellant system is shown schematically in Fig.10.
Figure 10: Basic Flow Scheme of Bipropellant Systems
T
P
P P
MON MMH
TT
Temperature Sensor
Tank for Pressurant Gas
Fill Valve (H )
Pressure Transducer
Pyro-Valve
(normally closed)
Filter
Pressure Regulator
Pressure Relief Valve
Check Valve
Pyro-Valve
(normally open)
Test Port
Temperature Sensor
Propellant Tank(surface tension)
Fill Valves (MON/MMH)
Pressure Transducer
Pyro-Valve(normally closed)
Filter
Thruster
e
HeT
P
P P
MON MMH
TT
Temperature Sensor
Tank for Pressurant Gas
Fill Valve (H )
Pressure Transducer
Pyro-Valve
(normally closed)
Filter
Pressure Regulator
Pressure Relief Valve
Check Valve
Pyro-Valve
(normally open)
Test Port
Temperature Sensor
Propellant Tank(surface tension)
Fill Valves (MON/MMH)
Pressure Transducer
Pyro-Valve(normally closed)
Filter
Thruster
e
He
- 34 -
The propellants are injected separately into the thruster combustion
chamber where they react spontaneously (hypergolic propellant) to
perform high-temperature, low molecular weight combustion
products, which are then expelled through a nozzle. A typical
bipropellant thruster configuration is shown schematically in Fig.11
below. Typical thrust range is 4 to 500N for spacecraft attitude and
orbit control.
Figure 11: Bipropellant Thruster Configuration
The Bipropellant system basically consists of a pressurising-gas
system, propellant tanks (with surface tension propellant
- 35 -
management devices), propellant lines and thrusters. Unlike
hydrazine thrusters, bipropellant thrusters accept only a limited range
of propellant inlet pressure variation of < 2. Therefore, the high-
pressure gas, generally helium, contained in a separate high pressure
tank, is regulated to the desired tank pressure, e.g. 17.5 bar. This
mode of operation is also referred to as the pressure constant mode.
The system contains check valves upstream of the propellant tanks to
prevent possible back-flow, mixing, and combustion of the
propellant vapours in the common pressurant gas line. Relief valves
are incorporated in the system upstream of the propellant tanks to
prevent system rupture in the event of a pressure regulator failure.
Filters are provided in the propellant lines upstream of the line valves
to prevent damage of the valve seats or clogging of injectors of
thrusters by contamination. Finally, the systems contains pyro- or
latch valves, line pressure transducers, fill and drain valves and
various test ports for system check out.
System-specific Impulse, Issp
In the case of the constant pressure mode, which is the common
mode of tank pressurisation for storable bipropellants, C is usually
close to 1 (e.g. 0.95) and the mass of the tank containing the
pressurising gas has to be added to the tankage mass; see Fig.10
above.
To include the constant pressure mode in our calculations, Eq. (37)
has to be modified to include the mass of the extra gas storage tank.
For the mass of the pressurising gas plus the extra gas storage tank
we get with Eq. (28) - as already derived for compressed gases:
MK
RTzmmm
pr
pr
prTprpr 1 (38)
- 36 -
The pressurising gas will have to fill the propellant tank plus the gas
storage tank at the end of the spacecraft mission. Therefore, with the
gas storage tank estimated to have a volume of about 10% of that of
the propellant tank, the mass of the pressurising gas can be calculated
with the help of the gas law (see Eq. (26)):
zRT
MVpm
Top
pr
1.1 (39)
With help of Eqs. (38) and (39), Eq. (36) can be now expanded to:
HWTprprpTHWPSS mmmmmmm =
xMK
RTz
zRT
Mp
zRT
MCK
KC
pm
pr
propP
P
op
p
11
1.1)1(11
(40)
With (21), (23) and (40) we obtain the final expression for the
system-spec. impulse of systems operating with liquid propellants in
the ”CONSTANT PRESSURE MODE”:
HWTprprpT
tot
ssp mmmmm
II
xMK
RTz
zRT
Mp
zRT
MCK
KC
p
v
pr
propP
P
op
e
111.1)1(
11
(41)
- 37 -
From Eq. (41) it is obvious that, as in the case of the blow-down
mode, both the type of propellant and propellant storage conditions
as well as the non-impulse dependent propulsion system mass x have
a major effect on the Issp. Because of the even more energetic bi-
propellant combinations, when compared with monopropellant
hydrazine, higher values of ve, and Issp can be achieved; see Table 6.
For comparison, a calculated Issp-value for typical propellant storage
and non-impulse system mass parameters are presented in Table 6,
showing an overall good agreement with Issp-values of actual built
spacecraft propulsion systems.
Table 6: Actual Bipropellant Propulsion Systems Performances (Listed data are examples and therefore only indicative)
PROPELLANT
THRUSTER
SPEC.-
IMPULSE
Isp (mission
average)
(Ns/kg)
TOTAL
IMPULSE
Itot
(Ns)
PROPUL-
SION
SYSTEM
MASS
mPS
(kg)
SYSTEM
SPEC.-
IMPULSE
Issp
(Ns/kg)
REMARKS
Actual Propulsion
Systems
Issp values derived
from Ref. 10
Bi-Propellant 2963 2.22 106 849 2615 DFS
Fuel:Monomethyl- 2900 2.89 106 1101 2625 EUROSTAR
2900 3.10 106 1170 2650 EUTELSAT-2 hydrazine (MMH) 2900 2.70 106 1147 2354 GALILEO CH3N2H3 2900 2.20 106 847 2597 INMARSAT-2 Oxidiser: Nitrogen 2930 5.05 106 1839 2746 OLYMPUS Tetroxide (NTO)
N2O4 2960 3.05 106 1147 2659 TVSAT/TDF1/
TELE-X
ve 3120 m/s 2900 3.34 106 1253 2666 TELECOM-2
r = 1.65; mix. ratio
1.15·103 kg/m
3
Max. op. pressure:
pop =17.5 bar
He pressurant gas:
K= 105 m
2/s
2
Kevlar tank
2950 - - 2639 Calculated: Issp for
surface tension tank:
C=0.95,
K=3.3104 m2/s2
x = 4.5% (large S/C)
- 38 -
Conclusion
Bipropellant systems are more complex and therefore more
expensive than monopropellant hydrazine systems. However, their
potential high system costs is compensated by their higher impulse
performance (high Issp) resulting in lower propulsion mass fraction
allowing a higher payload mass. Therefore, bipropellant systems are
mainly used for commercial spacecrafts with missions of high
impulse requirements. E.g. for geostationary communication
satellites, they form a single unified propulsion system, giving
maximum flexibility in the shared use of the propellant between the
orbit transfer operation, as well as the apogee and attitude control
functions.
ADVANTAGES:
- high Isp: ( 2900 Ns/kg) for F 25 N, Issp ( 2800 Ns/kg)
( 3110 Ns/kg) for F 500 N,
- high thrust capability, - up to 45 000 N.
DISADVANTAGES:
- system complexity with added valves, regulators, etc.;
- higher cost in comparison to monopropellant hydrazine systems.
- 39 -
4.2.5 General System Design Considerations
In order to ensure safety of personnel during spacecraft ground
operations, in general the following pressure ratings of
pressurized systems have to be followed:
- The burst pressure (causing rupture) of the
integrated system shall be not less than four
times the maximum system operating pressure.
Only the tank burst pressure in general is two
times the maximum propellant storage pressure
for the reason of low tank mass.
- The proof pressure (checking safety) of the
integrated system shall be not less than 1.5 times
the maximum system operating pressure. The
system has to pass successfully the proof
pressure before operating it for the first time.
This applies also for the case of system repair
where faulty equipment has to be replaced. After
repair, again the system has to pass successfully
a proof pressure cycle.
In general, propellant feed systems are an all-welded design in order
to minimize mass and ensure leak-tightness. Screw mounted
connections are used only for the connection of thrusters. This allows
easy mounting and even later replacement of this equipment if
required.
All components, which are in contact with the propellants are
designed for and have demonstrated their long term compatibility.
Therefore, high strength titanium alloy 6AL4V and pure titanium
A40 are normally used for tanks and all other components including
tubing lines.
- 40 -
4.2.6 Solid Propellant
The solid propellant rocket motor consists of a motor case,
containing a propellant grain, a nozzle and an igniter. The schematic
is shown in Fig.12.
There are two principal types of propellants:
- homogeneous propellants, which are composed of fuels that
contain enough chemically bonded oxygen to sustain the
propellant burning process,
- composite propellants, which are a mixture of powdered
metal (fuel), crystalline oxidiser and a polymer binder.
Most common is the use of composite propellants, usually based on
solid aluminium powder held in a hydroxyl terminated polybutadiene
(HTPB) synthetic rubber binder and stable solid oxidiser like
ammonium perchlorate (AP). The propellant is premixed and batch
loaded into lightweight simple motors.
Figure 12: Schematic of Solid Propellant Motor
SCHEMATIC OF SOLID PROPELLANT MOTOR
Igniter Motor Case Nozzle
Propellant Grain
Hot GasExhaust
- 41 -
System-specific Impulse, Issp
With regard to solid propellant rocket motors propulsion, Eq. (34)
of vaporising liquids (see Chapter 4.2.1 above) may be applied. In a
solid motor (see Fig.12), the propellant tank and the combustion
chamber are contained in the motor case. The motor case mcase mT
is filled with propellant mP according to the volumetric loading
fraction Ccase ( 90%), and during motor operation, the motor
chamber pressure will be: pc pop.
Therefore, with Eq. (34) the System-spec. Impulse becomes for
SOLID PROPELLANT ROCKET MOTORS:
xKC
p
v
mmm
II
casecase
c
e
HWcaseP
totssp
11
(42)
Table 7: Performances of Actual Solid Propellant Motor Systems (Listed data are examples and therefore only indicative)
PROPELLANT
THRUSTER
SPEC.-
IMPULSE
Isp (mission
average)
(Ns/kg)
TOTAL
IMPULSE
Itot
(Ns)
PROPUL-
SION
SYSTEM
MASS
mPS
(kg)
SYSTEM
SPEC.-
IMPULSE
Issp
(Ns/kg)
REMARKS
Actual Propulsion
Systems
Solid Propellant
(HTPB)
2842 7.73 106 2960 2611 Orbus-6 Inert.Upp.
Stage Motor [1]
ve ≤ 3000 m/s 2852 1.17 106 447 2617 MAGE 1S Apogee
Kick Motor [11]
2880 1.41 106 528 2670 MAGE 2 Apogee
Kick Motor [11]
2858 4.58 106 1729 2649 Solid End-burning
Motor [12]
2842
−
−
2620
Calculated: Issp for
Pc= 5.8 106 N/m2,
C=0.92, x=0%
K=4.2·104 m2/s2,
ρ=1.76·103 kg/m3
- 42 -
From Eq. (42) it is obvious, that the motor exhaust velocity ve ≡ Isp,
as well as the type of propellant (ρ), the motor case performance
factor Kcase, and the volumetric loading fraction Ccase influence the
Issp. As to be seen from Eq. (42), for high values of Issp, above all the
thrust exhaust velocity ve shall be high while the hardware mass x
shall be low. All other parameters are of secondary importance for
the value of the Issp. Table 7 shows a good overall agreement of
calculated with actual values of Issp.
Conclusion
In general, solid propulsion motors can only deliver their total
impulse potential in one firing, because off-modulation is not
possible. Therefore the usage of solid propulsion is restricted to:
- orbit change (e.g. apogee or perigee manoeuvre);
- impart acceleration (e.g. liquid reorientation maneuvers,
separation maneuvers).
ADVANTAGES:
- relatively simple operation;
- very high mass fraction, excellent bulk density and
packaging characteristics;
- good, long-term storage characteristics.
DISADVANTAGES:
- not readily tested and checked-out prior to flight;
- very difficult to stop and restart, throttle, pulse, etc.
(hybrid);
- limited Isp performance (2800 - 3000 Ns/kg);
- limited redundancy with associated reliability and safety
issues.
- 43 -
4.3 Electric Propulsion
4.3.1 Propulsion Concepts
In order to increase propulsion system impulse performance, e.g. for
interplanetary missions, the jet exhaust velocity has to be increased
beyond the ve ≤ 5000 m/s, which is best available from chemical
propulsion. This can be achieved by electrical propulsion that relies
on externally provided electric power to accelerate the working fluid
(propellant) to produce useful thrust. There are three main methods
by which the electrical energy may be converted into the kinetic
energy of thrust:
- Electrothermal Systems, where the propellant (gas) is
heated by passing it over an electric heated solid surface
(resistojet), or by passing it through an arc discharge
(arcjet). The heated gas is then accelerated by gas-dynamic
expansion in a nozzle. Typical applications of this principle
are the monopropellant hydrazine operated Power
Augmented Catalytic Thruster (PACT) and the Hydrazine-
Arcjet.
Figure 13: Schematic of Resistojet Thruster
Heater
Power Supply Gas Inlet Heat Exchanger
_
+Hot Gas
Exhaust
- 44 -
- Electromagnetic Systems, where a gas is heated in an arc
discharge to such a high temperature, that it is converted to
neutral plasma (plasma thruster). The plasma is then
expelled at high velocity by the interaction of the discharge
current with the magnetic field (Lorentz force). A typical
application of this principle is the Magneto-Plasma-
Dynamic (MPD) type of thruster.
Figure 14: Schematic of Magnetic Arcjet Thruster
- Electrostatic Systems, where usually a high molecular
propellant, such as xenon gas, is ionised (ion thruster) by
e.g. electron bombardment (Kaufman thrusters), or in a high
frequency electro-magnetic field (radio-frequency thrusters)
or by extracting ions from the surface of a liquid metal
(caesium) under the effect of a strong electrostatic field
(field emission). The ions are then accelerated to high
velocity (30 to 100 km/s) by a strong electric field.
Electrons are injected into the exhaust ion beam from an
electron emitter in order to keep it electrically neutral, thus
preventing an electric charge build-up of the spacecraft.
+
Gas Inlet Cathode AnodePower Supply
Plasma
Exhaust
Arc-Discharge
- 45 -
To the above described category of ion thrusters, the
Stationary Plasma Thruster (SPT), which belongs to the
category of Hall-effect Thrusters, uses an applied magnetic
field to control electrons in a quasi-neutral plasma
discharge.
Figure 15: Schematic of Ion Thruster
Finally, as an example of liquid propellants for electrostatic
electric propulsion, Caesium (Cs) is the propellant of choice
with a melting point of 29°C. Caesium, as a liquid metal, is
also desirable because it has a high atomic mass and
effectively wets metal surfaces. It is used for field emission
thrusters, or Field Emission Electric Propulsion (FEEP)
devices.
In a FEEP thruster, a strong electric field is established at
the tip (tailored cones) of a pair of closely spaced electrodes,
which even form a capillary propellant feed system for the
liquid caesium. See also the schematic of the FEEP thruster
which is shown below.
Ioniser Exit GridIoniserPropellant Supply
Positiv Ion Beam
Electron Emitter
Power Supply
AcceleratorGrid
+_
- 46 -
Fig. 16: FEEP Thruster Schematic
When the field reaches a threshold value, which is in the
order of 106 V/mm (for caesium), atoms on the surface of the
tip of the electrodes are ionised and eventually removed.
They are then accelerated to a high velocity in between the
positive emitter (tailored cones) and the negative accelerator
electrode. Expelled ions are replenished by the flow of liquid
propellant in the capillary feed system. A separate neutraliser
is required to maintain charge neutrality of the system.
FEEP-thrusters can achieve very high exhaust velocities up to
105 m/s at the expense of low thrust levels. This is due to the
limit of power available from the spacecraft. E.g., if we take
P = 1kW available, assuming an overall power conversion
efficiency of η=0.5, then with Eq. (13) we will get for the
thrust force: 010.010
5.021000
5
F N.
Conclusion: Thrust levels of electric propulsion are << thrust
levels of chemical propulsion
Field Emission Electric Propulsion (FEEP) Thruster
• Very High Isp
6000 to 10000 s.
• F = 10 N to 2 mN
• Cesium, Rubidium, Indium.
• Efficiency = 98%
- (Ion~30%; PPT~17
• Self contained propellant
reservoir.
• No moving parts.
- 47 -
4.3.2 Propulsion System Design and Performance
An electric propulsion system consists of a power generator (solar or
nuclear), power processing system (unit), electric control unit,
thruster assembly, propellant storage and propellant feed system (see
also Fig. 3, Chapter 4.1 above).
Figure 17: Electric Propulsion System Block Diagram
- Power Generator
Electric power can be obtained from either sunlight or from a
nuclear reactor. In the case of solar electric propulsion, solar
Plasma/Ion-Jet Power
Processing
Unit
Control Unit
Solar Array Solar Array
Propellant
Tank
Propellant
Feed
System
S/C
Interface
Gimbal
Thruster
- 48 -
photons are converted into electricity by solar cells. In nuclear
electric propulsion, thermal energy from the nuclear reactor is
converted into electricity by either a static or dynamic thermal-to-
electric power conversion system. Static, thermoelectric systems
have the advantage of no moving parts for high reliability, but they
have low efficiency. Dynamic systems have moving parts (e.g.,
turbines, generators, etc.) and do not scale well for small systems,
but they do have a higher efficiency.
- Power Processing System
Power processing systems are required to convert the voltage from
the power generator to the form required by the electric thruster.
For example, a solar array produces low-voltage DC (typically ~
100 V); this would need to be converted (via transformers, etc.) to
kilovolt levels for use in an ion thruster. The power processing
system is often referred to as the Power-Processing Unit (PPU).
- Electric Control Unit to operate electrically valves and thrusters.
- Propellant Storage & Feed Systems
Various combinations of propellant and thruster are possible,
depending on the specific application. In general, liquid or gaseous
propellants are stored and fed to the thruster assemblies as in
chemical propulsion. Details see also [1].
System-specific Impulse, Issp
Electric propulsion relies on externally provided electric power to
create or augment the kinetic energy of the exhaust jet. Therefore, for
the evaluation of the system-spec. impulse, the mass of the electric
power (supply- and processing) system, as depicted in Fig.3 and
- 49 -
Fig.17, has to be considered in addition to the propellant storage
system as already dealt with for chemical propulsion systems.
To describe the performance of electric propulsion systems, the
denominator of Eq. (22) has to be further determined. Both, the mass
of the propellant storage system and the mass of the electric power
supply system have to be considered.
For systems operation with gaseous propellants, e.g. xenon, the
combined mass of tank and propellant is calculated according to the
Eq. (28) as derived for cold gas systems above. The mass of the
electric power supply and processing system is calculated with the
system specific power W/kg:
PmEl (43)
where:
2ev
FP (44)
is the system input of electrical energy and = Pjet/P is the overall
energy conversion efficiency.
With Eqs. (29), (43) and (44) the combined mass of the propellant
storage system and the electric power system is calculated for
systems operating with gaseous propellants:
HWElTPHWElPSS mmmmmmm
xm
Fv
KM
zRTm
P
e
P
1
21
(45)
- 50 -
The system-specific impulse becomes with Eqs. (22) and (45):
xm
Fv
KM
zRT
v
mmmm
II
P
e
e
HWElTP
tot
ssp
12
1
(46)
And for:
ee
Pv
F
v
Ftm
(47)
with t =, which is the thruster operating time (s), the system-spec.
impulse for ELECTRIC PROPULSION SYSTEMS, operating with
gaseous propellants, becomes finally:
xv
KM
zRT
vI
e
e
ssp
12
12
(48)
Eq. (48) shows that the Issp for electric propulsion systems depends
on the parameters of propellant storage as well as on the energy
supply and processing systems. According to Eq. (48), for high
values of Issp, high values of ve and low values of x are required.
However, with regard to the impact of the energy supply and
processing system on the Issp, low values of ve and high values of γ
and as well as long thrust operation times τ are preferred. With
regard to the controversial requirement for ve, there must be optimal
values of ve-opt, which will result in maximum values of Issp.
- 51 -
The term of an optimal exhaust velocity ve-opt can be elucidated
schematically by the following picture:
Figure 18: Term of Optimal Exhaust Velocity
With increasing exhaust velocity ve, the combined mass of propellant
and tank is decreasing while the mass of the power supply (with
power processing) system is increasing. The point of inter-section of
the two curves determines the minimum of the system mass by ve-opt
resulting in a maximum value of Issp.
This diagram shows clearly, that with increasing thrusters exhaust
velocity ve, the mass of propellant becomes less important for the
mass of electric propulsion systems and therefore the system-spec.
impulse Issp, describes better the system performance than the usually
used Isp, which is only propellant related.
Maximum values of Issp can be derived by observing the first and
second derivates of Eq. (48) with regard to ve.
- 52 -
The first derivate of Eq. (48) with regard to ve has to be set equal to
zero:
0
1
)(2
b
va
v
dv
dI
dv
d
e
e
e
ssp
e
(49)
where:
KM
zRTa and 2b (50)
Explicitly it follows with Eq. (49)
2
2
2
1
1
0)(
b
va
b
va
Idv
d
e
e
ssp
e
(51)
Hence maximum and minimum values of Issp will be for:
)1( abve (52)
For a maximum value of Issp, the second derivate of Eq. (48) shall
be <0.
It follows for the second derivate of Eq. (48):
0
1
)1(2
)1(2
22
2
b
va
b
ab
abIvd
d
e
ssp
e
, (53)
- 53 -
resulting in a maximum value for Issp with the optimal thruster
exhaust velocity of:
KM
zRTv opte 12 (54)
In a similar way, the system-spec. impulse and the optimal thruster
exhaust velocity can be determined for electric propulsion systems
operating with liquid propellants, by taking into account the relevant
expressions for mass of the propellant storage systems as derived for
vaporising liquid gas- or hot gas systems above. For ELECTRIC
PROPULSION operating with e.g. vaporising liquid gas, the system-
spec. impulse becomes:
xv
KC
p
vI
eop
e
ssp
12
12
(55)
The Issp will be a maximum for the optimal thruster exhaust velocity
of:
KC
pv
op
opte
12 (56)
Therefore, for electric propulsion high impulse performance is not
dictated by maximum exhaust velocity, like for chemical propulsion,
but rather by optimum values of thrusters exhaust velocity ve-opt.
Here, high values of Issp, that is high values of ve-opt, will be achieved
mainly for high values of overall specific power , overall power
conversion efficiency , and thrust operation time . The thrust
operation time will be mainly dictated by mission manoeuvre
operating times and/or max. life of thrusters.
- 54 -
Parameters of the xenon gas storage system, like gas compressibility
factor z, tank performance factor K, gas storage temperature T, and
gas molecular mass M, will have only a secondary impact on values
of Issp.
A precise quantitative determination of the Issp of electric propulsion
systems is more difficult than for chemical propulsion systems. In the
case of electrical propulsion, the electrical power can be shared
partly and/or temporarily with the payload of a spacecraft. Here, Issp
is dependent on the operative conditions of a spacecraft. Therefore,
in order to allow a more quantitative comparison of actual electric
propulsion systems, Table 8 lists examples without considering their
power supply (solar array) systems. Listed examples are electrostatic
systems with Stationary Plasma Thrusters, SPT, and Kaufman-type
of ion thrusters.
Table 8: Performances of Actual Electric Propulsion Systems (Listed data are examples and therefore only indicative)
PROPELLANT
THRUSTER
SPEC.-
IMPULSE
Isp (mission
average)
(Ns/kg)
TOTAL
IMPULSE
Itot
(Ns)
PROPUL-
SION
SYSTEM
MASS
mPS
(kg)
SYSTEM
SPEC.-
IMPULSE
Issp
(Ns/kg)
REMARKS
Actual Propulsion
Systems
Electric
Propulsion Xenon
14700 7.65· 105 128 5980 GALS [13];
Stationary Plasma
Thruster, SPT-100
Xenon (Xe)
Mol.Mass (M):
15107 1.2106
111 10811 SMART-1; 14, 15;
PPS 1350 (SPT)
4 kg/kmol
21580 1.15·10
6 96 11980 ETS-VIII 16,
Kaufman-type Xenon
Ion Thruster
z = 0.3, at tank
pressure 150 bar;
K= 1105 m
2/s
2;
15000
−
−
11287
Calculated:
Xe-Propellant,
γ=82 W/kg; =50%,
=5000h, x=10%;
SPT-100
- 55 -
Table 8 notes the differences between the propulsion performance
reference numbers Isp (ve) and Issp. Hence, the differences becomes of
particular interest with respect to the calculation of the ‘propulsion
system mass fraction’, mSP/mSC. Usually this is done by taking into
account the rocket equation and calculating the mass of propellant,
Eq. (9). However, taking into account the entire propulsion system
mass, the ‘propulsion system mass fraction’ has to be calculated with
Eq. (19), which is related to the Issp.
The differences in calculating the ‘propulsion system mass fraction’
can be illustrated by the SMART-1 project (the first European
spacecraft travelled to and orbit the Moon [14] [15]) as noted in
Table 8. Results are elucidated in diagram Fig. 19 below.
0 1000 2000 3000 4000
0,00
0,05
0,10
0,15
0,20
0,25
0,30
0,35
0,40
(2)
(1)
(1) Calculated by System-spec. Impulse Equation
Issp
= 10 811 Ns/kg
ve = 15 107 m/s
(2) Calculated by Rocket Equation
Isp
= 15 107 Ns/kg
ve = 15 107 m/s
Mission to the Moon:
delta-v=3.7 km/s
(constant low thrust space manoeuvres)
DELTA-V PERFORMANCES OF S/C PROPULSION SYSTEMS
Calculated by System-spec. Impulse Equation versus Rocket Equation
mP
S/m
S/C
delta-v (m/s) Fig. 19: Delta-v Performance of Spacecraft Propulsion System
Diagram Fig.19 shows clearly that for electric propulsion, where the
electric subsystem mass, mEL, may form a major ‘dead’ dry mass of
the overall propulsion system mass, the ‘rocket equation’ does not
apply alone.
- 56 -
Table 9 presents a summary of the comparison between electrical
(ion) and chemical (bipropellant) propulsion.
Table 9: Comparison of typical electrical vs. chemical figures
Type of
Thruster
Spec. -
Impulse
(Ns/kg)
Thrust F
(N)
DC Power
Required
(W)
Electrical
(Ion thruster)
Chemical
(Bipropellant)
Order of
magnitude of
the ratio
ION/Chemical
30 000
≈ 3 000
101
10-3
– 0.2
4 – 500
10-4
400 – 800
4 – 8
(short term)
102
Conclusion:
- While chemical propulsion is limited to specific impulse exhaust
velocity of <5000 m/s, electric propulsion can achieve exhaust
velocities up to 100 000 m/s;
- Although electric propulsion results in very high specific impulse
Isp (ve), power is the major constraint for electric thrusters on
spacecraft. Therefore thrust force levels of electric propulsion will
be << thrust levels of chemical propulsion.
Main performance and operating characteristics of electric
propulsion are summarised together with chemical propulsion in
Tables 10 and 11, Chapter 5 below.
- 57 -
5 PROPULSION SYSTEMS SELECTION
CRITERIA
A detailed procedure for the selection of propulsion systems for
given spacecraft mission requirements is beyond the scope of this
booklet. The process for selecting and sizing the elements of
propulsion systems is detailed in 17.
However, in general, an important consideration for the selection of a
suitable propulsion system will be the trade-off between its impulse
or velocity increment (v) capability and the system mass.
Consequently, when selecting a spacecraft propulsion system for
given mission impulse demand, primarily the system will have to
meet the impulse or delta-v requirement with highest possible
spacecraft payload mass. Therefore, an important requirement of the
spacecraft designer will be that the mass of the propulsion system
shall be a minimum or at least shall not exceed a certain percentage
of the overall mass of the spacecraft. As already mentioned in
Chapter 3.2.2 above, the performance of propulsion systems cannot
be assessed only by the specific impulse Isp, but requires also taking
into account the system- specific impulse Issp.
To assess the suitability of spacecraft propulsion systems for
spacecraft mission impulse requirements, value ranges of Isp and Issp
(see Eq. (16)) of various actual built systems can be derived from
published data as summarised in Table 10. In addition, with Eq. (19)
the dependence of the propulsion system mass fraction mPS/mS/C, on
mission velocity increment v can be derived for any given value of
Isp (ve) and Issp. Curves of mPS/mS/C, plotted as a function of v for
different propulsion system designs with typical value ranges of Isp
and Issp from Table 10, are shown in diagram Fig. 20 below.
When suitable systems are selected, a refinement of the selection has
to be carried out. This process takes into consideration additional
parameters such as cost, operability, complexity and reliability, etc.
- 58 -
Table 10: Comparison of typical Spacecraft Propulsion Systems
Performances
Propellant Thruster-spec.
Impulse Isp
(mission average)
(Ns/kg)
System-spec.
Impulse, Issp
(Ns/kg)
Remarks
Cold Gas
Compressed Gas
(N2, A)
Vaporising
Liquid, (C3H8,
NH3)
510 – 706
618 – 800
193 – 283
486 – 654
Compressed Gas:
Titanium Tank
Liquid Gas:
Al-Tank with
Heat Exchanger
Liquid
Monopropellant
Hydrazine (N2H4)
2100 – 2 300
1440 -1860
Hot Gas; Tank
with Diaphragm
Bipropellant
MMH/ NTO,
(CH3N2H3/N2O4)
2900 – 3120
2354 - 2746 Hot Gas; Tank
with Surface
Tension Device
Solid (Composites,
HTPB)
2800 – 3000
2611 -2670 E.g. MAGE 1, 2
Apogee Kick
Motors
Electric
Propulsion
Electrostatic:
Stationary Plasma
Thruster (SPT)/
Kaufman-type Xe-
Ion Thruster
14000 - 34000
5980 - 11287
Xenon Propellant
E.g. GALS/
SMART-1/
ETS-VIII Projects
N.B.: Listed data, which can be derived from published data, are examples
and therefore only indicative.
- 59 -
500 1000 1500 2000 2500 3000
0,0
0,2
0,4
0,6
0,8
1,0
Issp
=5980 Ns/kg
NITROGEN (1
PROPANE
HYDRAZINE
PACT
SOLID PROPELLANT
BI_PROPELLANT
F1
F2
F3
Issp
=11980 Ns/kg
SPT/Ion Electric
Propulsion Systems
Bipropellant
Issp
=1440 Ns/kg
Hydrazine
Issp
=193 Ns/kg
(Nitrogen)
Issp
=2746 Ns/kg
Issp
=2356 Ns/kg
Issp
=1860 Ns/kg
Issp
=654 Ns/kg
(Ammonia)
Cold GasmP
S/m
S/C
delta-v (m/s)
Figure 20: Delta-V Performance Range of Spacecraft Propulsion System Concepts (Examples)
- 60 -
The diagram in Fig. 20 gives the first and most important indication
for the selection of propulsion systems. If we assume a system mass
ratio of mPS/mS/C < 0.30, we can read directly from Fig. 20:
- for low v 150 m/s, compressed cold gas and vaporising liquid
propulsion systems seem to be the best choice, because they meet
the requirement and have the lowest cost;
- for 150 v 650 m/s, monopropellant hydrazine fed propulsion
systems are the best choice, because of their inherent simplicity
(reliability) and potential low cost, while still meeting the Δv-
requirement;
- for high v 650 m/s, bipropellant systems, monopropellant
hydrazine fed resistojet systems (power-augmented thrusters,
arcjets), and electrostatic (electromagnetic) systems will satisfy the
v-requirements best.
Finally, for any given value of total impulse Itot, the mass of the
propulsion system mPS can be calculated of course directly from
values of Issp.
When a suitable spacecraft auxiliary propulsion system is selected,
however, the cost, complexity, operability and reliability of the
system also play an important role.
With regard to low v-requirements, compressed cold gas systems
used for auxiliary propulsion of spacecraft’s (attitude and orbit
control), although of moderate impulse capability, are still of interest
in view of their simplicity, high reliability, repeatability of impulse
bit and low system costs.
- 61 -
In considering the merits of the various compressed cold gas and
vaporising-liquid systems, the following major points must be
considered carefully:
- Additional heat may be necessary for vaporising, e.g.
propane for use in gas jets. For high thrust levels and
long thrust duration’s, this can give rise to thermal
problems in the propane system. As the latent heat of
ammonia is about three times that of propane, additional
technical problems may occur here.
- For zero-g conditions (non-spinning satellites), the
storage of liquefied propellants is more complex than that
of pressurised, inert gases, as tanks with bellows, surface
tension devices etc. have to be provided to separate liquid
and vapour. In addition, propellant gauging is much
more complex (with resulting higher costs) for liquefied
propellants under zero-g conditions than for compressed
gases. Moreover, fuel slosh of liquefied propellants may
cause extra problems for the dynamic behaviour of
spacecrafts.
Therefore, for low v-requirements, compressed cold gas systems
utilising N2 are the most commonly used.
For higher v-requirements, in the trade-off between monopropellant
hydrazine systems, bipropellant systems and electric propulsion
systems, the following major points have to be considered carefully:
- Because of their inherent simplicity, hydrazine
monopropellant systems still represent the lowest
possib1e cost technology in the field of liquid propulsion.
Such a technology is therefore of interest whenever a
moderate velocity increment is required or where mass is
not a critical design driver.
- 62 -
- For delivering low impulse-bits or impulses at low
spacecraft torques or acceleration forces, hydrazine
thrusters of potential low thrust levels (down to 0.5N)
will have to be used.
- High v-requirements for spacecraft in-orbit transfer and
attitude and orbit-control can only be met by bipropellant
systems (e.g. unified propulsion system) and electric
propulsion.
- For the selection of a hydrazine resistojet system (e.g.
power-augmented thrusters, arc-jet), thrust level and duty
cycle requirements have to be considered. The fact that
the limited power available and heat capacitance of the
electrothermal thruster impose a limit on thrust- and
duty-cycle levels may give rise to technical problems.
- With regard to electric propulsion like electromagnetic
and electrostatic systems, this technology, although still
under development, has proven to achieve thrusters
exhaust velocities ve an order of magnitude higher than
the best performing chemical propulsion systems; see
Table 9. Therefore, electric propulsion is essential for
further reduction of system (propellant) mass, enabling
higher payload mass and coping best with future high
energy mission requirements. However, depending on
thrust levels, electric propulsion can impose severe
power requirements on the spacecraft power supply and
power processing system.
When spacecraft propulsion systems are to be selected, the above
mentioned points have to be assessed properly with reference to the
flight mission and design requirements of the spacecraft itself.
Table 11 below presents an overall comparison summary of
candidate spacecraft propulsion systems.
- 63 -
Table 11: Survey: Typical Candidate Spacecraft Propulsion Systems
Type Thrust
Level
Range
(N)
Thruster
Exhaust
Velocity
(m/s)
Advantages Disadvantages
Cold Gas
(N2, A, NH3,
C3H8)
0.02 – 10
500 – 800
Extremely
simple, reliable,
very low cost
Very low
performance,
Solid Motor (e.g.
Apogee Kick
Motor)
28 000 –
47 000
2 800 –
3000
Simple, reliable,
relatively low
cost
Limited
performance, higher
thrust
Liquid:
Monopropellant
Hydrazine
(N2H4)
0.5 – 22
2 100 –
2 300
Simple, reliable,
low-cost
Moderate
performance
Bipropellant
(CH3N2H3/N2O4)
4 – 500
2 900 –
3 120
High
performance
More complicated
system than
monopropellant
Electric
Propulsion
Electrothermal:
Resistojet
(NH3, N2H4, H2)
5∙10-3
–
0.5
1 300 –
5 000
High
performance
Low thrust
Arcjet (NH3,
N2H4, H2, N2)
5∙10-2
– 5
4 000 –
15 000
High
performance,
High power,
complicated
interfaces
Electromagnetic:
Pulsed plasma,
(Teflon)
5∙10-6
–
5∙10-3
15 000
High
performance
High power, low
thrust, complicated
Electrostatic:
Stat. Plasma
Thruster (SPT)
(Xenon: Xe)
10-2
– 0.5
15 000 –
25 000
High
performance
High power, low
thrust, complicated
Ion (Hg, A, Xe) 10-3
– 0.2 30 000 Very high
performance
Very high power
N.B.: Listed data, which can be derived from published data, are examples
and therefore only indicative.
- 64 -
6 OUTLINE OF POTENTIAL FUTURE
SPACE PROPULSION
So far, chemical propulsion has given access to space and has even
taken spacecraft through the solar system. Electric propulsion, still
under development, offers a further vast increase in propulsion
system mass efficiency.
The prevailing goal of future propulsion in form of advanced
spacecraft propulsion systems is to enable cost efficient space
missions and extended exploration of the solar system up to
interstellar missions.
In order to achieve efficient mission costs, an important application
of advanced spacecraft propulsion is to reduce cost by:
- reduction of the total mass that must be launched from
Earth,
- reduction of propulsion system mass fraction, allowing
for higher payload mass,
- increase of mission impulse performance, allowing for
satellite extended orbit maintenance and attitude control.
A second goal of advanced spacecraft propulsion is to perform
extended (manned) exploration of the solar system and previously
‘impossible’ missions, like interstellar travel.
Consequently, the evolution of advanced spacecraft propulsion
systems will mainly focus on increased performance that is high
values of Issp.
In a first instance, advanced propulsion systems can be derived from
existing systems, by increasing the performance of chemical and
electric propulsion with regard to their mission impulse and velocity-
increment, Δv, capabilities.
- 65 -
6.1 Potential Improvement of Chemical
Propulsion
For chemical propulsion high performance, i.e. high values of Issp, resulting in low values of ‘propulsion system mass fraction’, is
primarily dictated by maximum values of Isp (ve). See to this Eq. (21)
with the mass of propellant storage system (propellant + tank), mPSS,
which is proportional to the system impulse capability and sized by
Isp (ve); see Eq. (10). However, the performance of state-of the-art
spacecraft engines operating with cold and hot gas can be considered
near to the theoretical limit for actual space storable propellant
combinations.
But the Issp can be still improved by also taking into account the non-
impulse system mass mH/W, - see Eq. (21). The emerging class of
micro-and nanospacecrafts requires miniaturization of the
propulsion system with help of ‘Microelectromechanical System’
(MEMS) technology for acceptable values of Issp, in order to achieve
a low value of mH/W.
Fig.21: MICRO PROPULSION: Laser Induced Etched Nozzle with
thrust force F= 0.5 ÷ 10 mN (Courtesy of Ångström Space
Technology Centre, Uppsala/Sweden)
- 66 -
For further reading about MEMS technology in particular with
reference to its space applications [18] is recommended.
In addition, with increasing interest in environmental and safety
issues, non-toxic monopropellant systems are under development.
Current satellite users and manufacturers are looking for more
environmentally friendly, safer propellants. Safer propellants can
reduce costs by eliminating the need for self-contained atmospheric
protective ensemble (SCAPE) suits that are needed for toxic
propellants by personnel for propellant filling and draining
operations. Also, extensive and prohibitive propellant safety
precautions and isolation of the space vehicle from parallel activities
during propellant loading operations can be minimised or eliminated.
Therefore, if environmentally safe and low toxic propellants are
used, the costs for operating satellites on ground can be lowered, in
some cases even dramatically.
A new family of environmentally friendly monopropellants has been
identified as an alternative to hydrazine. These new propellants are
based on blends of e.g. hydroxyl ammonium nitrate (HAN),
ammonium dinitramide (ADN), hydrazinium nitroformate (HNF),
nitrous oxide (N2O), and hydrogen peroxide (H2O2). When compared
to hydrazine, e.g. HAN and ADN blends have a range of specific
impulse (Isp) which can exceed that of hydrazine [19]. Testing of
HAN and ADN based propellants has begun to show promise and
could soon be adopted for spacecraft on-board propulsion systems
use.
To summarize, actual designs of chemical spacecraft propulsion
systems are well developed, but are being mainly complemented by
miniaturised cold/hot gas-, as well as by low-toxic monopropellant
systems.
- 67 -
6.2 Potential Improvement of Electric Propulsion
Most promising for further increase of propulsive performance
capabilities is the use of electric propulsion.
For electric propulsion, high values of Issp will be achieved mainly
for high values of ve-opt, which requires in particular high values of
overall specific power γ (watt per unit mass), combined with high
overall power conversion efficiency η, resulting in low values of mEl.
This can be illustrated by considering planetary missions to the edge
of the solar system, - mission to rendezvous with Pluto or other
members of the Kuiper belt. For constant low thrust maneuvers, a
total Δv ≥ 37 km/s has been assumed to cover escape from the
Earth’s gravitational field as well as escape from the solar system
with no gravity assist maneuvers, thus giving wide launch windows
[20]. A rough propulsion system analysis will show the needs for
further performance improvements.
Parametric investigations have been performed by altering overall
system specific power and thrust time , considering an overall
system power efficiency of = 0.67. In order to demonstrate the
impact of these parameters on Issp with resulting values of
‘propulsion system mass fraction’, parameters have been combined
for extreme cases of and as follows. Overall system values of containing electric power generators, power processing systems and
thrusters have been assumed in the range of = 30 ÷ 130 W/kg with
a main emphasis on Radioisotope Thermoelectric Generators (RTG),
needed for deep space missions. Performances of RTG’s have been
assumed for = 33 W/kg at 100 kW to = 625 W/kg for potential
future RTG designs at 10 MW [2]. Thruster operation times have
been assumed for max. life of thrusters, ranging from 7000 h to
20000 h [20]. In addition, it is assumed, that systems are operating
with xenon-gas propellant.
- 68 -
Parametric investigations performed according to Eqs. (55) with (56)
are depicted in Fig. 22 for optimal thruster exhaust velocity ve-opt.
0 10 000 20 000 30 000 40 000
0,0
0,1
0,2
0,3
0,4
0,5
0,6
0,7
0,8
0,9
1,0(1) (2)
(3)
(4)
Assumption: Constant low thrust maneuvers
Ion Propulsion, Xenon propellant
(4) Issp
= 54 499 Ns/kg
ve-opt
= 115070 m/s,
gamma = 130 W/kg,
tau = 20 000 h
(3) Issp
= 47 799 Ns/kg,
ve-opt
= 100 923 m/s,
gamma = 100 W/kg,
tau = 20 000 h
(2) Issp
= 28 278 Ns/kg,
ve-opt
= 59 707 m/s,
gamma = 100 W/kg,
tau = 7000 h
(1) Issp
= 15 489 Ns/kg,
ve-opt
= 32 703 m/s,
gamma = 30 W/kg,
tau = 7 000h
DELTA-V PERFORMANCES OF S/C ELECTRIC PROPULSION SYSTEMS
mP
S/m
S/C
delta-v (m/s) Fig. 22: Delta-v Performance of Potential Future Electric Propulsion
The parametric investigation performed by altering and shows the
importance of the large range of specific power mainly caused by
the electric power generators which have a major impact on the
overall value of . Here mainly power supply systems with high
specific power γ need to be further developed to achieve high values
of Issp, while for thrusters, the “Dual-Stage 4-Grid” type of gridded
ion thruster with a capability of ve ≤ 186 000 m/s is already under
development [20]. In addition, for high values of Issp, thrust operation
times should be always a maximum within the frame of mission
maneuver time in order to minimise power consumption, thus lower
mass of power supply system. This could imply multiple (cluster)
thruster configurations with thrusters operating serial in time in order
to obtain extra long thrust operation time, if required.
- 69 -
6.3 New Approaches
New approaches are studied or are under development, with
examples like:
- Solar- thermal Propulsion:
Fig. 23: Schematic of Solar-Thermal Propulsion System
(Courtesy by SNECMA)
Propulsion system is based on using solar radiation energy to heat
liquid hydrogen propellant which is passed through a heat
exchanger, reaching very high temperatures up to 2500 oK, before
expanding through a nozzle. The advantage would be:
higher thrust levels than achieved for electric propulsion, e.g.
F = 5 to 10 N continuous for 70 kW (solar),
higher exhaust velocities than achieved for chemical
propulsion, e.g. ve =8000 m/s.
Status: Several concepts for solar-thermal propulsion systems
have been proposed, however, so far none have been realised.
FUEL
thrust
chamber
Heat exchanger
reflector
FUEL
thrust
chamber
Heat exchanger
reflector
- 70 -
- Nuclear-thermal rockets:
There are two main different categories of nuclear technology
for space power and propulsion:
Radioisotope thermoelectric generators (RTG) and close-
cycle (e.g. Sterling technology) for nuclear electric power,
NEP, to power electric propulsion. Flight heritage of RTG’s
with power level < 10 kWe while future NEP’s aim at
>10 kWe to MWe’s for electric propulsion: ve = 20 000m/s to
≤ 186 000 m/s [20].
Open-cycle nuclear thermal reactors, NTR, which heat e.g.
liquid hydrogen propellant directly to produce rocket thrust.
Liquid hydrogen propellant absorbs heat from the core of a
fission reactor, before expanding through a nozzle: ve = 8000
to 9000 m/s, F = 20kN to 70 kN
Fig.24: Schematic of NTR
[Type a quote from the document or the summary of an interesting point. You can position the text box anywhere
in the document. Use the Text Box Tools tab to change the formatting of the pull quote text box.]
SCHEMATIC OF A NUCLEAR ROCKET ENGINE
H2 Propellant Turbin
Nuclear Reactor
Reactor Fuel Element
Nozzle Coolant
Jacket
- 71 -
Extensive research has been performed into nuclear-thermal
rockets in USA in 1960 as part of the NERVA program.
Status: Environmental and political concern about safe ground
test and launch of fueled reactor has reduced research in nuclear
technology.
- Exotic propulsion methods, such as:
Exotic Propulsion Systems are those “far out” ideas still under
study and are pure speculation so far. They will be required for
the ultimate dream of space exploration to travel to other star
systems, as depicted in TV shows like ‘Star Trek’. Two examples
of such exotic propulsion systems are outlined below.
Antimatter Propulsion: Matter- antimatter annihilation offers
the highest possible physical energy density of any known
reaction substance. Since matter and antimatter annihilate each
other completely, it is an incredibly compact way of storing
energy. E.g. a round trip to Mars with a 100-ton payload might
require only 30 gram of antimatter. However, sufficient
production and storage of antimatter (with potential complex
and high storage system mass) is still very much in the future.
Photon Propulsion: The generation of usable thrust by ejection
of photons is still very hypothetic. The generation of photons
by e.g. laser technology and their subsequent decay in space,
involves the mass-energy transfer expressed by Einstein’s
equation, E = mc2. Consequently, very large quantities of
energy will be required even for nominal levels of thrust.
Possibly, matter-antimatter annihilation can be harnessed for
photon propulsion in the future
For a further reading about advanced propulsion systems, [21] is
recommended.
- 72 -
7 GROUND TESTING OF PROPULSION
SYSTEMS
The following parameters can be measured during system ground
operations:
- pressure;
- temperature;
- electric current and voltage on e.g. solenoid valves.
The following tests can be performed:
- functional tests of electric actuated valves by measuring
electric current and voltages response time of valves;
- measurement of leak tightness of components, parts of
systems and overall systems by measurements of
temperatures and pressures at various points of time;
- measurement of leak tightness of the integrated propulsion
system with the help of a gas spectrometer, either by
“sniffing” or during test operations of the spacecraft in a
vacuum chamber;
- measurement of leak tightness of valves (thrusters) with
the help of 'glass pipettes'.
- 73 -
8 MISSION SURVEILLANCE OF
PROPULSION SYSTEMS
During spacecraft mission operations the following propulsion
system related telemetry data are available:
- pressure;
- temperature;
- thruster operations;
- system valves operations.
The propulsion systems can be checked for the following items:
- leak detection of systems (parts and overall system) by:
measurement of pressure and temperature of different
parts of systems at various points of time;
- propellant consumption and remaining mass of propellant by:
evaluation of pressure and temperature data (gas law) in
propellant tanks, related to initial mass of propellant at
beginning of life (BOL), also called ‘P.V.T’ (Pressure,
Volume, Temperature) method;
'book keeping' of thruster operations;
- thrust and spec. impulse of individual thrusters by:
evaluation of propellant consumption during in-orbit
thruster operations and comparison of planned and
achieved in orbit spacecraft movements.
- 74 -
9 LITERATURE/REFERENCES
1 G. P. Sutton, Oscar Biblarz, 2001, “Rocket Propulsion
Elements”, 7th Edition, John-Wiley & Sons, Ltd.,
ISBN: 0-471-32642-9,
2 P. Hill, R. Peterson, "Mechanics and Thermodynamics of
Propulsion". Addison-Wesley Publishing Company, Inc., USA.
ISBN 0-201-14659-2,
3 P. Erichsen, “Performance Evaluation of Spacecraft Propulsion
Systems in Relation to Mission Impulse Requirements”, ESA,
1997, SP-398, Proceedings of the Second European Spacecraft
Propulsion Conference, 27-29 May, 1997,
4 P. Erichsen, “A Quick-Look Analysis Tool for the Impulse
Performance of Spacecraft Propulsion Systems”, EUCASS,
Europe, 2007, EUCASS Paper 01-05-03,
5 D.R. Trotsenburg, “A Design Tool for Low Thrust Rocket
Propulsion Systems”, MSc. Thesis in Aerospace Engineering
TU-Delft, August 2004,
6 Adib Najib, „Ermittlung der impulsunabhängigen Masse im
Verhältnis zu den impulsabhängigen Massen für Raumfahrt-
antriebssysteme“, Study work at the Institute for Aerospace,
Technical University Berlin, Jan. 2004,
[7 Lee B. Holcomb; “Satellite Auxiliary-Propulsion Selection
Techniques”, NASA Technical Report 32-1505, November 1,
1970,
[8 COS-B Project, AOCS Design Specification, Ref. D-310.0200,
dated 15 March 1975,
- 75 -
[9 W. Inden; “Development Results of the ESRO TD Satellite
Pneumatic System”, Paper Reprint from Lecture Series No. 45
on Attitude Stabilization of Satellites in Orbit, AGARD,
10 D. Gale, et. al., “Bibliography of Liquid Propellant Propulsion
Systems (LPPS) of European Spacecraft”, Study Note:
SN/ESA-P/001/89/BAe Issue 1, January 1990,
11 P. Erichsen, “Catalogue of Propulsion Motors for Spacecraft”,
ESTEC Working Paper No. 1348, September 1982,
[12] H.F.R. Schöyer, “Some New European Developments in
Chemical Propulsion”, ESA Bulletin No. 66, 1991,
[13] A. Bober et al., “Development and Qualification Test of a SPT
Electric Propulsion System for ‘GALS’ Spacecraft”; IEPC-93-
008, IEPC-93-008, 23rd
Int. Electric Prop. Conf. 1993,
[14] D. Estublier et. al., “Electric Propulsion on SMART-1”, ESA
Bulletin 129, February 2007,
[15] J. Kugelberg, “Accommodating Electric Propulsion on a Small
Spacecraft”, IAF-00.S.4.09, 2000,
[16] T. Ozaki et. al., “In orbit Operation of 20mN Class Xenon Ion
Engine for ETS-VIII”, IEPC-2007-084, 30th
Int. Electric Prop.
Conf. 2007,
17 W. J. Larsson and J. R. Wertz, “Space Mission Analysis and
Design”, Chapter 17: Space Propulsion Systems, Microcosm,
Inc. Torrance, California and Kluwer Academic Publishers
Dodrecht/Boston/London,
- 76 -
18 J. Köhler, “Bringing Silicon Microsystems to Space”, Acta
Universitatis Upsaliensis, Uppsala 2001,
19] M. Sc Niklas Boman, Dr M. Ford, ”Reduced Hazard Propellant
– Propulsion System Impact”, Proc. ’2nd
Int. Conference on
Green Propellants for Space Propulsion’, Cagliari, Sardinia,
Italy, 7-8 June 2004 (ESA SP-557, October 2004),
20] D.G. Fearn, R. Walker, “Interstellar Precursor Missions using
Advanced Dual-Stage Ion Propulsion Systems”, International
Workshop on Innovative System Concepts. ESTEC, 21
February 2006,
21 M. Tajmar, “Advanced Space Propulsion Systems”, Springer-
Verlag/Wien, 2003 (ISBN 3-211-83862-7).
Notes