Hydrogen (1H) has three naturally occurring isotopes: 1H, 2H, and 3H. 1H and 2H are stable, while 3H has a half-life of 12.32 years. Heavier isotopes also exist; all are synthetic and have a half-life of less than 1 zeptosecond (10−21 s).
Hydrogen is the only element whose isotopes have different names that remain in common use today: 2H is deuterium and 3H is tritium. The symbols D and T are sometimes used for deuterium and tritium; IUPAC (International Union of Pure and Applied Chemistry) accepts said symbols, but recommends the standard isotopic symbols 2H and 3H, to avoid confusion in alphabetic sorting of chemical formulas. 1H, with no neutrons, may be called protium to disambiguate. (During the early study of radioactivity, some other heavy radioisotopes were given names, but such names are rarely used today.)
The three most stable isotopes of hydrogen: protium (A = 1), deuterium (A = 2), and tritium (A = 3)
List of isotopes
Note: "y" means year, but "ys" means yoctosecond (10−24 second).
1H (atomic mass 1.007825031898(14) Da) is the most common hydrogen isotope, with an abundance of > 99.98%. Its nucleus consists of only a single proton, so it has the formal name protium.
The proton has never been observed to decay, so 1H is considered stable. It is the only stable nuclide with no neutrons. Some Grand Unified Theories proposed in the 1970s predict that proton decay can occur with a half-life between 1028 and 1036 years. If so, then 1H (and all nuclei now believed to be stable) are only observationally stable. As of 2018, experiments have shown that the mean lifetime of the proton is > 3.6×1029 years.
Deuterium consists of 1 proton, 1 neutron, and 1 electron.
Deuterium, 2H (atomic mass 2.014101777844(15) Da), the other stable hydrogen isotope, has one proton and one neutron in its nucleus, called a deuteron. 2H
comprises 26–184 ppm (by population, not mass) of hydrogen on Earth;
the lower number tends to be found in hydrogen gas and higher enrichment
(150 ppm) is typical of seawater. Deuterium on Earth has been enriched with respect to its initial concentration in the Big Bang and outer Solar System (≈ 27 ppm, atom fraction) and older parts of the Milky Way (≈ 23 ppm). Presumably the differential concentration of deuterium in the inner Solar System is due to the lower volatility of deuterium gas and compounds, enriching deuterium fractions in comets and planets exposed to significant heat from the Sun over billions of years of Solar System evolution.
Deuterium is not radioactive, and is not a significant toxicity hazard. Water enriched in 2H is called heavy water. Deuterium and its compounds are used as a non-radioactive label in chemical experiments and in solvents for 1H-nuclear magnetic resonance spectroscopy. Heavy water is used as a neutron moderator and coolant for nuclear reactors. Deuterium is also a potential fuel for commercial nuclear fusion.
Tritium can be used in chemical and biological labeling experiments as a radioactive tracer. Deuterium–tritium fusion uses 2H and 3H as its main reactants, giving energy through the loss of mass when the two nuclei collide and fuse at high temperatures.
Hydrogen-4
4H (atomic mass4.02643(11) Da), with one proton and three neutrons, is a highly unstable isotope. It has been synthesized in the laboratory by bombarding tritium with fast-moving deuterons; the triton captured a neutron from the deuteron. The presence of 4H was deduced by detecting the emitted protons. It decays by neutron emission into 3H with a half-life of 139(10) ys (1.39(10)×10−22 s).
In the 1955 satirical novel The Mouse That Roared, the name quadium was given to the 4H that powered the Q-bomb that the Duchy of Grand Fenwick captured from the United States.
Hydrogen-5
5H (atomic mass5.03531(10) Da),
with one proton and four neutrons, is highly unstable. It has been
synthesized in the lab by bombarding tritium with fast-moving tritons; one triton captures two neutrons from the other, becoming a nucleus
with one proton and four neutrons. The remaining proton may be detected,
and the existence of 5H deduced. It decays by double neutron emission into 3H and has a half-life of 86(6) ys (8.6(6)×10−23 s) – the shortest half-life of any known nuclide.
Hydrogen-6
6H (atomic mass6.04496(27) Da) has one proton and five neutrons. It has a half-life of 294(67) ys (2.94(67)×10−22 s). In 2025, 6H was produced using an 855 MeV electron beam impinging upon on a 7Li target.
Hydrogen-7
7H (atomic mass7.05275(108) Da) has one proton and six neutrons. It was first synthesized in 2003 by a group of Russian, Japanese and French scientists at Riken's Radioactive Isotope Beam Factory by bombarding hydrogen with helium-8
atoms; all six of the helium-8's neutrons were donated to the hydrogen
nucleus. The two remaining protons were detected by the "Riken
telescope", a device made of several layers of sensors, positioned
behind the target of the RI Beam cyclotron. 7H has a half-life of 652(558) ys (6.52(558)×10−22 s).
Decay chains
4H and 5H decay directly to 3H, which then decays to stable 3He. Decay of the heaviest isotopes, 6H and 7H, has not been experimentally observed.
Decay times are in yoctoseconds (10−24 s) for all these isotopes except 3H, which is in years.
( ) – Uncertainty (1σ) is given in concise form in parentheses after the corresponding last digits.
Carvings of fauna common in the Sahara during the wet phase, found at Tassili in the central Sahara
The Sahara pump theory is a hypothesis that explains how flora and fauna migrated between Eurasia and Africa via a land bridge in the Levant region (the Levantine corridor). It posits that extended periods of abundant rainfall lasting many thousands of years (pluvial periods) in Africa are associated with a "wet-green Sahara" phase, during which larger lakes and more rivers existed (see North African climate cycles).
This caused changes in the flora and fauna found in the area. Migration
along the river corridor was halted when, during a desert phase
1.8–0.8 million years ago (mya), the Nile ceased to flow completely and possibly flowed only temporarily in other periods due to the geologic uplift (Nubian Swell) of the Nile River region.
Mechanism
During periods of a wet or Green Sahara, the Sahara and Arabia become a savanna grassland and African flora and fauna become common. Following inter-pluvial arid periods, the Sahara area then reverts to
desert conditions, usually as a result of the retreat of the West African Monsoon southwards. Evaporation exceeds precipitation, the level of water in lakes like Lake Chad falls, and rivers become dry wadis. Flora and fauna previously widespread as a result retreat northwards to the Atlas Mountains, southwards into West Africa, or eastwards into the Nile Valley and thence either southeast to the Ethiopian Highlands and Kenya or northeast across the Sinai into Asia. This separates populations of some of the species in areas with different climates, forcing them to adapt, possibly giving rise to allopatric speciation.
Plio-Pleistocene
The Plio-Pleistocene migrations to Africa included the Caprinae in two waves at 3.2 Ma and 2.7–2.5 Ma; Nyctereutes at 2.5 Ma, and Equus at 2.3 Ma. Hippotragus migrated at 2.6 Ma from Africa to the Siwaliks of the Himalayas. Asian bovids moved to Europe and to and from Africa. The primate Theropithecus experienced contraction and its fossils are found only in Europe and Asia, while Homo and Macaca settled wide ranges.
185,000–20,000 years ago
Between
about 133 and 122 thousand years ago (kya), the southern parts of the
Saharan-Arabian Desert experienced the start of the Abbassia Pluvial,
a wet period with increased monsoonal precipitation, around
100-200 mm/year. This allowed Eurasian biota to travel to Africa and
vice versa. The growth of speleothems
(which requires rainwater) was detected in Hol-Zakh, Ashalim, Even-Sid,
Ma'ale-ha-Meyshar, Ktora Cracks, Nagev Tzavoa Cave. In Qafzeh and Es
Skuhl caves, where at that time precipitation was 600–1000 mm/year, the
remains of Qafzeh-Skhul type anatomically modern humans are dated from this period, but human occupation seems to end in the later arid period.
The Red Sea
coastal route was extremely arid before 140 and after 115 kya. Slightly
wetter conditions appear at 90–87 kya, but it still was just one tenth
the rainfall around 125 kya. Speleothems are detected only in Even-Sid-2.
In the southern Negev Desert speleothems did not grow between 185–140 kya (MIS 6), 110–90 (MIS 5.4–5.2), nor after 85 kya nor during most of the interglacial period (MIS 5.1), the glacial period and Holocene. This suggests that the southern Negev was arid to hyper-arid in these periods.
The coastal route around the western Mediterranean
may have been open at times during the last glacial; speleothems grew
in Hol-Zakh and in Nagev Tzavoa Caves. Comparison of speleothem
formation with calcite horizons suggests that the wet periods were
limited to only tens or hundreds of years.
From 60–30 kya there were extremely dry conditions in many parts of Africa.
Last Glacial Maximum
An example of the Saharan pump has occurred after the Last Glacial Maximum
(LGM). During the Last Glacial Maximum the Sahara desert was more
extensive than it is now with the extent of the tropical forests being
greatly reduced. During this period, the lower temperatures reduced the strength of the Hadley cell whereby rising tropical air of the Intertropical Convergence Zone (ITCZ) brings rain to the tropics, while dry descending air, at about 20 degrees north,
flows back to the equator and brings desert conditions to this region.
This phase is associated with high rates of wind-blown mineral dust,
found in marine cores that come from the north tropical Atlantic.
Around 12,500 BC, the amount of dust in the cores in the Bølling–Allerød phase suddenly plummets and shows a period of much wetter conditions in the Sahara, indicating a Dansgaard–Oeschger
(DO) event (a sudden warming followed by a slower cooling of the
climate). The moister Saharan conditions had begun about 12,500 BC, with
the extension of the ITCZ northward in the northern hemisphere summer,
bringing moist wet conditions and a savanna climate to the Sahara, which
(apart from a short dry spell associated with the Younger Dryas) peaked during the Holocene thermal maximum
climatic phase at 4000 BC when mid-latitude temperatures seem to have
been between 2 and 3 degrees warmer than in the recent past. Analysis
of Nile River deposited sediments in the delta also shows this period had a higher proportion of sediments coming from the Blue Nile, suggesting higher rainfall also in the Ethiopian Highlands. This was caused principally by a stronger monsoonal circulation throughout the sub-tropical regions, affecting India, Arabia and the Sahara. Lake Victoria only recently became the source of the White Nile and dried out almost completely around 15 kya.
The sudden subsequent movement of the ITCZ southwards with a Heinrich event (a sudden cooling followed by a slower warming), linked to changes with the El Niño–Southern Oscillation
cycle, led to a rapid drying out of the Saharan and Arabian regions,
which quickly became desert. This is linked to a marked decline in the
scale of the Nile floods between 2700 and 2100 BC. One theory proposed that humans accelerated the drying out period from 6,000–2,500 BC by pastoralists overgrazing available grassland.
Middle Paleolithic: Homo heidelbergensis into the Middle East and Western Europe.
Upper Paleolithic: Homo sapiens (possible early "Out of Africa" wave, receded before 80,000 years ago and eventually replaced by the "coastal migration" wave after 70,000 years ago)
Rocket propellant is used as a reaction mass ejected 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 rear-ward, 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 (fuel) must be present in the 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
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. The 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. For future planetary missions the use of local resources and
solar energy for in situ propellant production is considered.
Solid chemical propellants
Solid propellants come in two main types. "Composites" are mostly 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 such as RDX, HMX, aluminium, beryllium. Plasticizers, stabilizers, and burn rate modifiers (iron oxide, copper oxide) can be added.
Single-, double-, or triple-bases are homogeneous mixtures of one
to three primary ingredients, which must include fuel and oxidizer, and
often include binders and plasticizers. All components are
macroscopically indistinguishable and often blended as liquids and cured
in a single batch. Ingredients can have multiple roles: RDX is both
fuel and oxidizer, while nitrocellulose is fuel, oxidizer, and
structural polymer.
Many propellants contain elements of double-base and composite
propellants, and often contain 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 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.
Extraterrestrial on-site production is being explored by combining aluminum and ice (ALICE).
Advantages
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 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
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 stage 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 ensure that 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.
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-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.
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 launch vehicle.
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 9, Falcon Heavy, Soyuz, Zenit, Angara, and Long March 6,
among others. 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.
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
proportional to 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.[10]
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 7 times 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.
Hybrid propellants consist of a storable liquid oxidizer used with a
solid fuel, which retains most virtues of both liquids (high ISP) and solids (simplicity).
A hybrid-propellant 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 that depend on
the rocket's acceleration to transport the fluid into the combustion
chamber. Fewer fluids typically mean fewer and smaller piping systems,
valves, and pumps.
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 the 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), a 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 oxidizer flux can lead to flooding and loss of
flame-holding, which 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 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 of
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 crewed 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.
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.
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.
For low-performance applications, such as attitude control jets, compressed 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.
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