A Delta IV Heavy during liftoff. The rocket is entirely fuelled with Liquid Hydrogen and Liquid Oxygen cryogenic propellants.
Rocket propellant is the reaction mass of a rocket. This reaction mass is ejected at the highest achievable velocity from a rocket engine to produce thrust. The energy required can either come from the propellants themselves, as with a chemical rocket, or from an external source, as with ion engines.
Overview
Rockets create thrust by expelling mass backward at high velocity. The thrust produced can be calculated by multiplying the mass flow rate of the propellants by their exhaust velocity relative to the rocket (specific impulse). A rocket can be thought of as being accelerated by the pressure of the combusting gases against the combustion chamber and nozzle, not by "pushing" against the air behind or below it. Rocket engines perform best in outer space
because of the lack of air pressure on the outside of the engine. In
space it is also possible to fit a longer nozzle without suffering from
flow separation.
Most chemical propellants release energy through redox chemistry, more specifically combustion. As such, both an oxidizing agent and a reducing agent must be present in the fuel mixture. Decomposition, such as that of highly unstable peroxide bonds in monopropellant rockets, can also be the source of energy.
In the case of bipropellant liquid rockets, a mixture of reducing fuel and oxidizing oxidizer is introduced into a combustion chamber, typically using a turbopump to overcome the pressure. As combustion takes place, the liquid propellant mass
is converted into a huge volume of gas at high temperature and
pressure. This exhaust stream is ejected from the engine nozzle at high
velocity, creating an opposing force that propels the rocket forward in
accordance with Newton's laws of motion.
Chemical rockets can be grouped by phase. Solid rockets use propellant in the solid phase, liquid fuel rockets use propellant in the liquid phase, gas fuel rockets use propellant in the gas phase, and hybrid rockets use a combination of solid and liquid or gaseous propellants.
In the case of solid rocket motors, the fuel and oxidizer are
combined when the motor is cast. Propellant combustion occurs inside the
motor casing, which must contain the pressures developed. Solid rockets
are typically have higher thrust, less specific impulse, shorter burn times, and a higher mass than liquid rockets, and additionally cannot be stopped once lit.
Rocket stages
In space, the maximum change in velocity that a rocket stage can impart on its payload is primarily a function of its mass ratio and its exhaust velocity. This relationship is described by the rocket equation. Exhaust velocity is dependent on the propellant and engine used and closely related to specific impulse,
the total energy delivered to the rocket vehicle per unit of propellant
mass consumed. Mass ratio can also be affected by the choice of a given
propellant.
Rocket stages that fly through the atmosphere usually use lower
performing, high molecular mass, high-density propellants due to the
smaller and lighter tankage required. Upper stages, which mostly or only
operate in the vacuum of space, tend to use the high energy, high
performance, low density liquid hydrogen fuel.
Solid chemical rockets
Solid propellants come in two main types. "Composites" are composed mostly of a mixture of granules of solid oxidizer, such as ammonium nitrate, ammonium dinitramide, ammonium perchlorate, or potassium nitrate in a polymer binding agent, with flakes or powders of energetic fuel compounds (examples: RDX, HMX, aluminium, beryllium). Plasticizers, stabilizers, and/or burn rate modifiers (iron oxide, copper oxide) can also be added.
Single-, double-, or triple-bases (depending on the number of
primary ingredients) are homogeneous mixtures of one to three primary
ingredients. These primary ingredients must include fuel and oxidizer
and often also include binders and plasticizers. All components are
macroscopically indistinguishable and often blended as liquids and cured
in a single batch. Ingredients can often have multiple roles. For
example, RDX is both a fuel and oxidizer while nitrocellulose is a fuel,
oxidizer, and plasticizer.
Further complicating categorization, there are many propellants
that contain elements of double-base and composite propellants, which
often contain some amount of energetic additives homogeneously mixed
into the binder. In the case of gunpowder (a pressed composite without a
polymeric binder) the fuel is charcoal, the oxidizer is potassium
nitrate, and sulphur serves as a reaction catalyst while also being
consumed to form a variety of reaction products such as potassium sulfide.
The newest nitramine solid propellants based on CL-20 (HNIW) can match the performance of NTO/UDMH storable liquid propellants, but cannot be throttled or restarted.
Advantages of solid propellants
Solid
propellant rockets are much easier to store and handle than liquid
propellant rockets. High propellant density makes for compact size as
well. These features plus simplicity and low cost make solid propellant
rockets ideal for military applications.
Their simplicity also makes solid rockets a good choice whenever
large amounts of thrust are needed and the cost is an issue. The Space Shuttle and many other orbital launch vehicles use solid-fueled rockets in their boost stages (solid rocket boosters) for this reason.
Disadvantages of solid propellants
Solid fuel rockets have lower specific impulse,
a measure of propellant efficiency, than liquid fuel rockets. As a
result, the overall performance of solid upper stages is less than
liquid stages even though the solid mass ratios are usually in the .91
to .93 range, as good as or better than most liquid propellant upper
stages. The high mass ratios possible with these unsegmented solid upper
stages is a result of high propellant density and very high
strength-to-weight ratio filament-wound motor casings.
A drawback to solid rockets is that they cannot be throttled in
real time, although a programmed thrust schedule can be created by
adjusting the interior propellant geometry. Solid rockets can be vented
to extinguish combustion or reverse thrust as a means of controlling
range or accommodating warhead separation. Casting large amounts of
propellant requires consistency and repeatability to avoid cracks and
voids in the completed motor. The blending and casting take place under
computer control in a vacuum, and the propellant blend is spread thin
and scanned to assure no large gas bubbles are introduced into the
motor.
Solid fuel rockets are intolerant to cracks and voids and require
post-processing such as X-ray scans to identify faults. The combustion
process is dependent on the surface area of the fuel. Voids and cracks
represent local increases in burning surface area, increasing the local
temperature, which increases the local rate of combustion. This positive
feedback loop can easily lead to catastrophic failure of the case or
nozzle.
Solid propellant history
Solid rocket propellant was first developed during the 13th century under the Chinese Song dynasty. The Song Chinese first used gunpowder in 1232 during the military siege of Kaifeng.
During the 1950s and 60s researchers in the United States developed ammonium perchlorate composite propellant (APCP). This mixture is typically 69-70% finely ground ammonium perchlorate (an oxidizer), combined with 16-20% fine aluminium powder (a fuel), held together in a base of 11-14% polybutadiene acrylonitrile (PBAN) or Hydroxyl-terminated polybutadiene
(polybutadiene rubber fuel). The mixture is formed as a thickened
liquid and then cast into the correct shape and cured into a firm but
flexible load-bearing solid. Historically the tally
of APCP solid propellants is relatively small. The military, however,
uses a wide variety of different types of solid propellants some of
which exceed the performance of APCP. A comparison of the highest
specific impulses achieved with the various solid and liquid propellant
combinations used in current launch vehicles is given in the article on solid-fuel rockets.
In the 1970s and 1980s, the U.S. switched entirely to solid-fueled ICBMs: the LGM-30 Minuteman and LG-118A Peacekeeper (MX). In the 1980s and 1990s, the USSR/Russia also deployed solid-fueled ICBMs (RT-23, RT-2PM, and RT-2UTTH), but retains two liquid-fueled ICBMs (R-36 and UR-100N).
All solid-fueled ICBMs on both sides had three initial solid stages,
and those with multiple independently targeted warheads had a precision
maneuverable bus used to fine tune the trajectory of the re-entry
vehicles.
Liquid chemical rocket propellants
The main types of liquid propellants are storable propellants, which tend to be hypergolic, and cryogenic propellants.
Advantages of liquid propellant
Liquid-fueled rockets have higher specific impulse
than solid rockets and are capable of being throttled, shut down, and
restarted. Only the combustion chamber of a liquid-fueled rocket needs
to withstand high combustion pressures and temperatures. Cooling can be
done regeneratively with the liquid propellant. On vehicles employing turbopumps,
the propellant tanks are at a lower pressure than the combustion
chamber, decreasing tank mass. For these reasons, most orbital launch
vehicles use liquid propellants.
The primary specific impulse advantage of liquid propellants is
due to the availability of high-performance oxidizers. Several
practical liquid oxidizers (liquid oxygen, dinitrogen tetroxide, and hydrogen peroxide) are available which have better specific impulse than the ammonium perchlorate used in most solid rockets when paired with suitable fuels.
Some gases, notably oxygen and nitrogen, may be able to be collected from the upper atmosphere, and transferred up to low-Earth orbit for use in propellant depots at substantially reduced cost.
Disadvantages of liquid propellants
The main difficulties with liquid propellants are also with the oxidizers. Storable oxidizers, such as nitric acid and nitrogen tetroxide,
tend to be extremely toxic and highly reactive, while cryogenic
propellants by definition must be stored at low temperature and can also
have reactivity/toxicity issues. Liquid oxygen (LOX) is the only flown cryogenic oxidizer - others such as FLOX, a fluorine/LOX mix, have never been flown due to instability, toxicity, and explosivity. Several other unstable, energetic, and toxic oxidizers have been proposed: liquid ozone (O3), ClF3, and ClF5.
Liquid-fueled rockets require potentially troublesome valves,
seals, and turbopumps, which increase the cost of the rocket. Turbopumps
are particularly troublesome due to high performance requirements.
Current cryogenic types
- Liquid oxygen (LOX) and highly refined kerosene (RP-1). Used for the first stages of the Atlas V, Falcon, Soyuz, Zenit, and developmental rockets like Angara and Long March 6. This combination is widely regarded as the most practical for boosters that lift off at ground level and therefore must operate at full atmospheric pressure.
- LOX and liquid hydrogen. Used on the Centaur upper stage, the Delta IV rocket, the H-IIA rocket, most stages of the European Ariane 5, and the Space launch system core and upper stages.
- LOX and liquid methane will be used on several rockets in development, including Vulcan, New Glenn, and SpaceX Starship.
Current storable types
- Dinitrogen tetroxide (N2O4) and hydrazine (N2H4), MMH, or UDMH. Used in military, orbital, and deep space rockets because both liquids are storable for long periods at reasonable temperatures and pressures. N2O4/UDMH is the main fuel for the Proton rocket, older Long March rockets (LM 1-4), PSLV, Fregat, and Briz-M upper stages. This combination is hypergolic, making for attractively simple ignition sequences. The major inconvenience is that these propellants are highly toxic and require careful handling.
- Monopropellants such as hydrogen peroxide, hydrazine, and nitrous oxide are primarily used for attitude control and spacecraft station-keeping where their long-term storability, simplicity of use, and ability to provide the tiny impulses needed outweighs their lower specific impulse as compared to bipropellants. Hydrogen peroxide is also used to drive the turbopumps on the first stage of the Soyuz launch vehicle.
Mixture ratio
The
theoretical exhaust velocity of a given propellant chemistry is
a function of the energy released per unit of propellant mass (specific
energy). In chemical rockets, unburned fuel or oxidizer represents the
loss of chemical potential energy, which reduces the specific energy.
However, most rockets run fuel-rich mixtures, which result in lower
theoretical exhaust velocities.
However, fuel-rich mixtures also have lower molecular weight
exhaust species. The nozzle of the rocket converts the thermal energy
of the propellants into directed kinetic energy. This conversion happens
in the time it takes for the propellants to flow from the combustion
chamber through the engine throat and out the nozzle, usually on the
order of one millisecond. Molecules store thermal energy in rotation,
vibration, and translation, of which only the latter can easily be used
to add energy to the rocket stage. Molecules with fewer atoms (like CO
and H2) have fewer available vibrational and rotational modes than molecules with more atoms (like CO2 and H2O).
Consequently, smaller molecules store less vibrational and rotational
energy for a given amount of heat input, resulting in more translation
energy being available to be converted to kinetic energy. The resulting
improvement in nozzle efficiency is large enough that real rocket
engines improve their actual exhaust velocity by running rich mixtures
with somewhat lower theoretical exhaust velocities.
The effect of exhaust molecular weight on nozzle efficiency is
most important for nozzles operating near sea level. High expansion
rockets operating in a vacuum see a much smaller effect, and so are run
less rich.
LOX/hydrocarbon rockets are run slightly rich (O/F mass ratio of 3 rather than stoichiometric
of 3.4 to 4) because the energy release per unit mass drops off quickly
as the mixture ratio deviates from stoichiometric. LOX/LH2
rockets are run very rich (O/F mass ratio of 4 rather than
stoichiometric 8) because hydrogen is so light that the energy release
per unit mass of propellant drops very slowly with extra hydrogen. In
fact, LOX/LH2 rockets are generally limited in how rich they
run by the performance penalty of the mass of the extra hydrogen tankage
instead of the underlying chemistry.
Another reason for running rich is that off-stoichiometric
mixtures burn cooler than stoichiometric mixtures, which makes engine
cooling easier. Because fuel-rich combustion products are less
chemically reactive (corrosive)
than oxidizer-rich combustion products, a vast majority of rocket
engines are designed to run fuel-rich. At least one exception exists:
the Russian RD-180 preburner, which burns LOX and RP-1 at a ratio of 2.72.
Additionally, mixture ratios can be dynamic during launch. This
can be exploited with designs that adjust the oxidizer to fuel ratio
(along with overall thrust) throughout a flight to maximize overall
system performance. For instance, during lift-off thrust is more
valuable than specific impulse, and careful adjustment of the O/F ratio
may allow higher thrust levels. Once the rocket is away from the
launchpad, the engine O/F ratio can be tuned for higher efficiency.
Propellant density
Although liquid hydrogen gives a high Isp,
its low density is a disadvantage: hydrogen occupies about 7x more
volume per kilogram than dense fuels such as kerosene. The fuel tankage,
plumbing, and pump must be correspondingly larger. This increases the
vehicle's dry mass, reducing performance. Liquid hydrogen is also
relatively expensive to produce and store, and causes difficulties with
design, manufacture, and operation of the vehicle. However, liquid
hydrogen is extremely well suited to upper stage use where Isp is at a premium and thrust to weight ratios are less relevant.
Dense propellant launch vehicles have a higher takeoff mass due to lower Isp,
but can more easily develop high takeoff thrusts due to the reduced
volume of engine components. This means that vehicles with dense-fueled
booster stages reach orbit earlier, minimizing losses due to gravity drag and reducing the effective delta-v requirement.
The proposed tripropellant rocket uses mainly dense fuel while at low altitude and switches across to hydrogen at higher altitude. Studies in the 1960s proposed single stage to orbit vehicles using this technique. The Space Shuttle
approximated this by using dense solid rocket boosters for the majority
of the thrust during the first 120 seconds. The main engines burned a
fuel-rich hydrogen and oxygen mixture, operating continuously throughout
the launch but providing the majority of thrust at higher altitudes
after SRB burnout.
Other chemical propellants
Hybrid propellants
Hybrid propellants: a storable oxidizer used with a solid fuel, which
retains most virtues of both liquids (high ISP) and solids
(simplicity).
A hybrid rocket
usually has a solid fuel and a liquid or NEMA oxidizer. The fluid
oxidizer can make it possible to throttle and restart the motor just
like a liquid-fueled rocket. Hybrid rockets can also be environmentally
safer than solid rockets since some high-performance solid-phase
oxidizers contain chlorine (specifically composites with ammonium
perchlorate), versus the more benign liquid oxygen or nitrous oxide
often used in hybrids. This is only true for specific hybrid systems.
There have been hybrids which have used chlorine or fluorine compounds
as oxidizers and hazardous materials such as beryllium compounds mixed
into the solid fuel grain. Because just one constituent is a fluid,
hybrids can be simpler than liquid rockets depending motive force used
to transport the fluid into the combustion chamber. Fewer fluids
typically mean fewer and smaller piping systems, valves and pumps (if
utilized).
Hybrid motors suffer two major drawbacks. The first, shared with
solid rocket motors, is that the casing around the fuel grain must be
built to withstand full combustion pressure and often extreme
temperatures as well. However, modern composite structures handle this
problem well, and when used with nitrous oxide
and a solid rubber propellant (HTPB), relatively small percentage of
fuel is needed anyway, so the combustion chamber is not especially
large.
The primary remaining difficulty with hybrids is with mixing the
propellants during the combustion process. In solid propellants, the
oxidizer and fuel are mixed in a factory in carefully controlled
conditions. Liquid propellants are generally mixed by the injector at
the top of the combustion chamber, which directs many small swift-moving
streams of fuel and oxidizer into one another. Liquid-fueled rocket
injector design has been studied at great length and still resists
reliable performance prediction. In a hybrid motor, the mixing happens
at the melting or evaporating surface of the fuel. The mixing is not a
well-controlled process and generally, quite a lot of propellant is left
unburned,
which limits the efficiency of the motor. The combustion rate of the
fuel is largely determined by the oxidizer flux and exposed fuel surface
area. This combustion rate is not usually sufficient for high power
operations such as boost stages unless the surface area or oxidizer flux
is high. Too high of oxidizer flux can lead to flooding and loss of
flame holding that locally extinguishes the combustion. Surface area can
be increased, typically by longer grains or multiple ports, but this
can increase combustion chamber size, reduce grain strength and/or
reduce volumetric loading. Additionally, as the burn continues, the hole
down the center of the grain (the 'port') widens and the mixture ratio
tends to become more oxidizer rich.
There has been much less development of hybrid motors than solid
and liquid motors. For military use, ease of handling and maintenance
have driven the use of solid rockets. For orbital work, liquid fuels
are more efficient than hybrids and most development has concentrated
there. There has recently been an increase in hybrid motor development
for nonmilitary suborbital work:
- Several universities have recently experimented with hybrid rockets. Brigham Young University, the University of Utah and Utah State University launched a student-designed rocket called Unity IV in 1995 which burned the solid fuel hydroxy-terminated polybutadiene (HTPB) with an oxidizer of gaseous oxygen, and in 2003 launched a larger version which burned HTPB with nitrous oxide. Stanford University researches nitrous-oxide/paraffin wax hybrid motors. UCLA has launched hybrid rockets through an undergraduate student group since 2009 using HTPB.
- The Rochester Institute of Technology was building an HTPB hybrid rocket to launch small payloads into space and to several near-Earth objects. Its first launch was in the Summer of 2007.
- Scaled Composites SpaceShipOne, the first private manned spacecraft, was powered by a hybrid rocket burning HTPB with nitrous oxide: RocketMotorOne. The hybrid rocket engine was manufactured by SpaceDev. SpaceDev partially based its motors on experimental data collected from the testing of AMROC's (American Rocket Company) motors at NASA's Stennis Space Center's E1 test stand.
Gaseous propellants
GOX (gaseous oxygen) was used as the oxidizer for the Buran program's orbital maneuvering system.
Inert propellants
Some rocket designs impart energy to their propellants with external energy sources. For example, water rockets use a compressed gas, typically air, to force the water reaction mass out of the rocket.
Ion thruster
Ion thrusters ionize a neutral gas and create thrust by accelerating the ions (or the plasma) by electric and/or magnetic fields.
Thermal rockets
Thermal rockets use inert propellants of low molecular weight that are chemically compatible with the heating mechanism at high temperatures. Solar thermal rockets and nuclear thermal rockets typically propose to use liquid hydrogen for a specific impulse
of around 600–900 seconds, or in some cases water that is exhausted as
steam for a specific impulse of about 190 seconds. Nuclear thermal
rockets use the heat of nuclear fission
to add energy to the propellant. Some designs separate the nuclear fuel
and working fluid, minimizing the potential for radioactive
contamination, but nuclear fuel loss was a persistent problem during
real-world testing programs. Solar thermal rockets use concentrated
sunlight to heat a propellant, rather than using a nuclear reactor.
Compressed gas
For low performance applications, such as attitude control jets, compressed inert gases such as nitrogen have been employed.
Energy is stored in the pressure of the inert gas. However, due to the
low density of all practical gases and high mass of the pressure vessel
required to contain it, compressed gases see little current use.
Nuclear plasma
In Project Orion and other nuclear pulse propulsion proposals, the propellant would be plasma debris from a series of nuclear explosions.
