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Text Content

ROCKET PROPELLANTS



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 * Liquids
 * Solids
 * Hybrids
 * Tables of Properties



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Propellant is the chemical mixture burned to produce thrust in rockets and
consists of a fuel and an oxidizer. A fuel is a substance that burns when
combined with oxygen producing gas for propulsion. An oxidizer is an agent that
releases oxygen for combination with a fuel. The ratio of oxidizer to fuel is
called the mixture ratio. Propellants are classified according to their state -
liquid, solid, or hybrid.

The gauge for rating the efficiency of rocket propellants is specific impulse,
stated in seconds. Specific impulse indicates how many pounds (or kilograms) of
thrust are obtained by the consumption of one pound (or kilogram) of propellant
in one second. Specific impulse is characteristic of the type of propellant,
however, its exact value will vary to some extent with the operating conditions
and design of the rocket engine.

Liquid Propellants

In a liquid propellant rocket, the fuel and oxidizer are stored in separate
tanks, and are fed through a system of pipes, valves, and turbopumps to a
combustion chamber where they are combined and burned to produce thrust. Liquid
propellant engines are more complex than their solid propellant counterparts,
however, they offer several advantages. By controlling the flow of propellant to
the combustion chamber, the engine can be throttled, stopped, or restarted.

A good liquid propellant is one with a high specific impulse or, stated another
way, one with a high speed of exhaust gas ejection. This implies a high
combustion temperature and exhaust gases with small molecular weights. However,
there is another important factor that must be taken into consideration: the
density of the propellant. Using low-density propellants means that larger
storage tanks will be required, thus increasing the mass of the launch vehicle.
Storage temperature is also important. A propellant with a low storage
temperature, i.e. a cryogenic, will require thermal insulation, thus further
increasing the mass of the launcher. The toxicity of the propellant is likewise
important. Safety hazards exist when handling, transporting, and storing highly
toxic compounds. Also, some propellants are very corrosive; however, materials
that are resistant to certain propellants have been identified for use in rocket
construction.

Liquid propellants used in rocketry can be classified into three types:
petroleum, cryogens, and hypergols.

Petroleum fuels are those refined from crude oil and are a mixture of complex
hydrocarbons, i.e. organic compounds containing only carbon and hydrogen. The
petroleum used as rocket fuel is a type of highly refined kerosene, called RP-1
in the United States. Petroleum fuels are usually used in combination with
liquid oxygen as the oxidizer. Kerosene delivers a specific impulse considerably
less than cryogenic fuels, but it is generally better than hypergolic
propellants.

Specifications for RP-1 where first issued in the United States in 1957 when the
need for a clean burning petroleum rocket fuel was recognized. Prior
experimentation with jet fuels produced tarry residue in the engine cooling
passages and excessive soot, coke and other deposits in the gas generator. Even
with the new specifications, kerosene-burning engines still produce enough
residues that their operational lifetimes are limited.

Liquid oxygen and RP-1 are used as the propellant in the first-stage boosters of
the Atlas and Delta II launch vehicles. It also powered the first-stages of the
Saturn 1B and Saturn V rockets.

Cryogenic propellants are liquefied gases stored at very low temperatures, most
frequently liquid hydrogen (LH2) as the fuel and liquid oxygen (LO2 or LOX) as
the oxidizer. Hydrogen remains liquid at temperatures of -253 oC (-423 oF) and
oxygen remains in a liquid state at temperatures of -183 oC (-297 oF).

Because of the low temperatures of cryogenic propellants, they are difficult to
store over long periods of time. For this reason, they are less desirable for
use in military rockets that must be kept launch ready for months at a time.
Furthermore, liquid hydrogen has a very low density (0.071 g/ml) and, therefore,
requires a storage volume many times greater than other fuels. Despite these
drawbacks, the high efficiency of liquid oxygen/liquid hydrogen makes these
problems worth coping with when reaction time and storability are not too
critical. Liquid hydrogen delivers a specific impulse about 30%-40% higher than
most other rocket fuels.

Liquid oxygen and liquid hydrogen are used as the propellant in the high
efficiency main engines of the Space Shuttle. LOX/LH2 also powered the upper
stages of the Saturn V and Saturn 1B rockets, as well as the Centaur upper
stage, the United States' first LOX/LH2 rocket (1962).

Another cryogenic fuel with desirable properties for space propulsion systems is
liquid methane (-162 oC). When burned with liquid oxygen, methane is higher
performing than state-of-the-art storable propellants but without the volume
increase common with LOX/LH2 systems, which results in an overall lower vehicle
mass as compared to common hypergolic propellants. LOX/methane is also clean
burning and non-toxic. Future missions to Mars will likely use methane fuel
because it can be manufactured partly from Martian in-situ resources.
LOX/methane has no flight history and very limited ground-test history.

Liquid fluorine (-188 oC) burning engines have also been developed and fired
successfully. Fluorine is not only extremely toxic; it is a super-oxidizer that
reacts, usually violently, with almost everything except nitrogen, the lighter
noble gases, and substances that have already been fluorinated. Despite these
drawbacks, fluorine produces very impressive engine performance. It can also be
mixed with liquid oxygen to improve the performance of LOX-burning engines; the
resulting mixture is called FLOX. Because of fluorine's high toxicity, it has
been largely abandoned by most space-faring nations.

Some fluorine containing compounds, such as chlorine pentafluoride, have also
been considered for use as an 'oxidizer' in deep-space applications.

Hypergolic propellants are fuels and oxidizers that ignite spontaneously on
contact with each other and require no ignition source. The easy start and
restart capability of hypergols make them ideal for spacecraft maneuvering
systems. Also, since hypergols remain liquid at normal temperatures, they do not
pose the storage problems of cryogenic propellants. Hypergols are highly toxic
and must be handled with extreme care.

Hypergolic fuels commonly include hydrazine, monomethyl hydrazine (MMH) and
unsymmetrical dimethyl hydrazine (UDMH). Hydrazine gives the best performance as
a rocket fuel, but it has a high freezing point and is too unstable for use as a
coolant. MMH is more stable and gives the best performance when freezing point
is an issue, such as spacecraft propulsion applications. UDMH has the lowest
freezing point and has enough thermal stability to be used in large
regeneratively cooled engines. Consequently, UDMH is often used in launch
vehicle applications even though it is the least efficient of the hydrazine
derivatives. Also commonly used are blended fuels, such as Aerozine 50 (or
"50-50"), which is a mixture of 50% UDMH and 50% hydrazine. Aerozine 50 is
almost as stable as UDMH and provides better performance.

The oxidizer is usually nitrogen tetroxide (NTO) or nitric acid. In the United
States, the nitric acid formulation most commonly used is type III-A, called
inhibited red-fuming nitric acid (IRFNA), which consists of HNO3 + 14% N2O4 +
1.5-2.5% H2O + 0.6% HF (added as a corrosion inhibitor). Nitrogen tetroxide is
less corrosive than nitric acid and provides better performance, but it has a
higher freezing point. Consequently, nitrogen tetroxide is usually the oxidizer
of choice when freezing point is not an issue, however, the freezing point can
be lowered with the introduction nitric oxide. The resulting oxidizer is called
mixed oxides of nitrogen (MON). The number included in the description, e.g.
MON-3 or MON-25, indicates the percentage of nitric oxide by weight. While pure
nitrogen tetroxide has a freezing point of about -9 oC, the freezing point of
MON-3 is -15 oC and that of MON-25 is -55 oC.

USA military specifications for IRFNA were first published in 1954, followed in
1955 with UDMH specifications.

The Titan family of launch vehicles and the second stage of the Delta II rocket
use NTO/Aerozine 50 propellant. NTO/MMH is used in the orbital maneuvering
system (OMS) and reaction control system (RCS) of the Space Shuttle orbiter.
IRFNA/UDMH is often used in tactical missiles such as the US Army's Lance
(1972-91).

Hydrazine is also frequently used as a monopropellant in catalytic decomposition
engines. In these engines, a liquid fuel decomposes into hot gas in the presence
of a catalyst. The decomposition of hydrazine produces temperatures up to about
1,100 oC (2,000 oF) and a specific impulse of about 230 or 240 seconds.
Hydrazine decomposes to either hydrogen and nitrogen, or ammonia and nitrogen.

Other propellants have also been used, a few of which deserve mentioning:

Alcohols were commonly used as fuels during the early years of rocketry. The
German V-2 missile, as well as the USA Redstone, burned LOX and ethyl alcohol
(ethanol), diluted with water to reduce combustion chamber temperature. However,
as more efficient fuels where developed, alcohols fell into general disuse.

Hydrogen peroxide once attracted considerable attention as an oxidizer and was
used in Britain's Black Arrow rocket. In high concentrations, hydrogen peroxide
is called high-test peroxide (HTP). The performance and density of HTP is close
to that of nitric acid, and it is far less toxic and corrosive; however it has a
poor freezing point and is unstable. Although HTP never made it as an oxidizer
in large bi-propellant applications, it has found widespread use as a
monopropellant. In the presence of a catalyst, HTP decomposes into oxygen and
superheated steam and produces a specific impulse of about 150 s.

Nitrous oxide has been used as both an oxidizer and as a monopropellant. It is
the oxidizer of choice for many hybrid rocket designs and has been used
frequently in amateur high-powered rocketry. In the presence of a catalyst,
nitrous oxide will decompose exothermically into nitrogen and oxygen and produce
a specific impulse of about 170 s.

Solid Propellants

Solid propellant motors are the simplest of all rocket designs. They consist of
a casing, usually steel, filled with a mixture of solid compounds (fuel and
oxidizer) that burn at a rapid rate, expelling hot gases from a nozzle to
produce thrust. When ignited, a solid propellant burns from the center out
towards the sides of the casing. The shape of the center channel determines the
rate and pattern of the burn, thus providing a means to control thrust. Unlike
liquid propellant engines, solid propellant motors cannot be shut down. Once
ignited, they will burn until all the propellant is exhausted.

There are two families of solids propellants: homogeneous and composite. Both
types are dense, stable at ordinary temperatures, and easily storable.

Homogeneous propellants are either simple base or double base. A simple base
propellant consists of a single compound, usually nitrocellulose, which has both
an oxidation capacity and a reduction capacity. Double base propellants usually
consist of nitrocellulose and nitroglycerine, to which a plasticiser is added.
Homogeneous propellants do not usually have specific impulses greater than about
210 seconds under normal conditions. Their main asset is that they do not
produce traceable fumes and are, therefore, commonly used in tactical weapons.
They are also often used to perform subsidiary functions such as jettisoning
spent parts or separating one stage from another.

Modern composite propellants are heterogeneous powders (mixtures) that use a
crystallized or finely ground mineral salt as an oxidizer, often ammonium
perchlorate, which constitutes between 60% and 90% of the mass of the
propellant. The fuel itself is generally aluminum. The propellant is held
together by a polymeric binder, usually polyurethane or polybutadienes, which is
also consumed as fuel. Additional compounds are sometimes included, such as a
catalyst to help increase the burning rate, or other agents to make the powder
easier to manufacture. The final product is rubber like substance with the
consistency of a hard rubber eraser.

Composite propellants are often identified by the type of polymeric binder used.
The two most common binders are polybutadiene acrylic acid acrylonitrile (PBAN)
and hydroxy-terminator polybutadiene (HTPB). PBAN formulations give a slightly
higher specific impulse, density, and burn rate than equivalent formulations
using HTPB. However, PBAN propellant is the more difficult to mix and process
and requires an elevated curing temperature. HTPB binder is stronger and more
flexible than PBAN binder. Both PBAN and HTPB formulations result in propellants
that deliver excellent performance, have good mechanical properties, and offer
potentially long burn times.

Solid propellant motors have a variety of uses. Small solids often power the
final stage of a launch vehicle, or attach to payloads to boost them to higher
orbits. Medium solids such as the Payload Assist Module (PAM) and the Inertial
Upper Stage (IUS) provide the added boost to place satellites into
geosynchronous orbit or on planetary trajectories.

The Titan, Delta, and Space Shuttle launch vehicles use strap-on solid
propellant rockets to provide added thrust at liftoff. The Space Shuttle uses
the largest solid rocket motors ever built and flown. Each booster contains
500,000 kg (1,100,000 pounds) of propellant and can produce up to 14,680,000
Newtons (3,300,000 pounds) of thrust.

Hybrid Propellants

Hybrid propellant engines represent an intermediate group between solid and
liquid propellant engines. One of the substances is solid, usually the fuel,
while the other, usually the oxidizer, is liquid. The liquid is injected into
the solid, whose fuel reservoir also serves as the combustion chamber. The main
advantage of such engines is that they have high performance, similar to that of
solid propellants, but the combustion can be moderated, stopped, or even
restarted. It is difficult to make use of this concept for vary large thrusts,
and thus, hybrid propellant engines are rarely built.

A hybrid engine burning nitrous oxide as the liquid oxidizer and HTPB rubber as
the solid fuel powered the vehicle SpaceShipOne, which won the Ansari X-Prize.






PROPERTIES OF ROCKET PROPELLANTS
  CompoundChemical
FormulaMolecular
WeightDensityMelting
PointBoiling
Point Liquid OxygenO232.001.14 g/ml-218.8oC-183.0oC Liquid FluorineF238.001.50
g/ml-219.6oC-188.1oC Nitrogen TetroxideN2O492.011.45 g/ml-9.3oC21.15oC Nitric
AcidHNO363.011.55 g/ml-41.6oC83oC Hydrogen PeroxideH2O234.021.44
g/ml-0.4oC150.2oC Nitrous OxideN2O44.011.22 g/ml-90.8oC-88.5oC Chlorine
PentafluorideClF5130.451.9 g/ml-103oC-13.1oC Ammonium
PerchlorateNH4ClO4117.491.95 g/ml240oCN/A Liquid HydrogenH22.0160.071
g/ml-259.3oC-252.9oC Liquid MethaneCH416.040.423 g/ml-182.5oC-161.6oC Ethyl
AlcoholC2H5OH46.070.789 g/ml-114.1oC78.2oC n-Dodecane
(Kerosene)C12H26170.340.749 g/ml-9.6oC216.3oC RP-1CnH1.953n≈1750.820
g/mlN/A177-274oC HydrazineN2H432.051.004 g/ml1.4oC113.5oC Methyl
HydrazineCH3NHNH246.070.866 g/ml-52.4oC87.5oC Dimethyl
Hydrazine(CH3)2NNH260.100.791 g/ml-58oC63.9oC AluminumAl26.982.70
g/ml660.4oC2467oC Polybutadiene(C4H6)n≈3000≈0.93 g/mlN/AN/A



NOTES: Chemically, kerosene is a mixture of hydrocarbons; the chemical
composition depends on its source, but it usually consists of about ten
different hydrocarbons, each containing from 10 to 16 carbon atoms per molecule;
the constituents include n-dodecane, alkyl benzenes, and naphthalene and its
derivatives. Kerosene is usually represented by the single compound n-dodecane.
RP-1 is a special type of kerosene covered by Military Specification
MIL-R-25576. In Russia, similar specifications were developed under
specifications T-1 and RG-1. Nitrogen tetroxide and nitric acid are hypergolic
with hydrazine, MMH and UDMH. Oxygen is not hypergolic with any commonly used
fuel. Ammonium perchlorate decomposes, rather than melts, at a temperature of
about 240 oC.




ROCKET PROPELLANT PERFORMANCE
  Combustion chamber pressure, Pc = 68 atm (1000 PSI) ... Nozzle exit pressure,
Pe = 1 atm OxidizerFuelHypergolicMixture RatioSpecific Impulse
(s, sea level)Density Impulse
(kg-s/l, S.L.) Liquid Oxygen Liquid HydrogenNo5.00381124 Liquid
MethaneNo2.77299235 Ethanol + 25% waterNo1.29269264 KeroseneNo2.29289294
HydrazineNo0.74303321 MMHNo1.15300298 UDMHNo1.38297286 50-50No1.06300300 Liquid
Fluorine Liquid HydrogenYes6.00400155 HydrazineYes1.82338432 FLOX-70
KeroseneYes3.80320385 Nitrogen Tetroxide KeroseneNo3.53267330
HydrazineYes1.08286342 MMHYes1.73280325 UDMHYes2.10277316 50-50Yes1.59280326
Red-Fuming Nitric Acid
(14% N2O4) KeroseneNo4.42256335 HydrazineYes1.28276341 MMHYes2.13269328
UDMHYes2.60266321 50-50Yes1.94270329 Hydrogen Peroxide
(85% concentration) KeroseneNo7.84258324 HydrazineYes2.15269328 Nitrous Oxide
HTPB (solid)No6.48248290 Chlorine Pentafluoride HydrazineYes2.12297439 Ammonium
Perchlorate
(solid) Aluminum + HTPB (a)No2.12277474 Aluminum + PBAN (b)No2.33277476





NOTES: Specific impulses are theoretical maximum assuming 100% efficiency;
actual performance will be less. All mixture ratios are optimum for the
operating pressures indicated, unless otherwise noted. LO2/LH2 and LF2/LH2
mixture ratios are higher than optimum to improve density impulse. FLOX-70 is a
mixture of 70% liquid fluorine and 30% liquid oxygen. Where kerosene is
indicated, the calculations are based on n-dodecane. Solid propellant
formulation (a): 68% AP + 18% Al + 14% HTPB. Solid propellant formulation (b):
70% AP + 16% Al + 12% PBAN + 2% epoxy curing agent.




SELECTED ROCKETS AND THEIR PROPELLANTS
  RocketStageEnginesPropellantSpecific Impulse Atlas/Centaur (1962)0
1
2Rocketdyne YLR89-NA7 (x2)
Rocketdyne YLR105-NA7
P&W RL-10A-3-3 (x2)LOX/RP-1
LOX/RP-1
LOX/LH2259s sl / 292s vac
220s sl / 309s vac
444s vacuum Titan II (1964)1
2Aerojet LR-87-AJ-5 (x2)
Aerojet LR-91-AJ-5NTO/Aerozine 50
NTO/Aerozine 50259s sl / 285s vac
312s vacuum Saturn V (1967)1
2
3Rocketdyne F-1 (x5)
Rocketdyne J-2 (x5)
Rocketdyne J-2LOX/RP-1
LOX/LH2
LOX/LH2265s sl / 304s vac
424s vacuum
424s vacuum Space Shuttle (1981)0
1
OMS
RCSThiokol SRB (x2)
Rocketdyne SSME (x3)
Aerojet OMS (x2)
Kaiser Marquardt R-40 & R-1EPBAN Solid
LOX/LH2
NTO/MMH
NTO/MMH242s sl / 268s vac
363s sl / 453s vac
313s vacuum
280s vacuum Delta II (1989)0
1
2Castor 4A (x9)
Rocketdyne RS-27
Aerojet AJ10-118KHTPB Solid
LOX/RP-1
NTO/Aerozine 50238s sl / 266s vac
264s sl / 295s vac
320s vacuum





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Compiled, edited and written in part by Robert A. Braeunig, 1996, 2005, 2006,
2008.
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