Geography of Mars

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/Geography_of_Mars

 

High-resolution colorized map of Mars based on Viking orbiter images. Surface frost and water ice fog brighten the impact basin Hellas to the right of lower center; Syrtis Major just above it is darkened by winds that sweep dust off its basaltic surface. Residual north and south polar ice caps are shown at upper and lower right as they appear in early summer and at minimum size, respectively.

 

The geography of Mars, also known as areography, entails the delineation and characterization of regions on Mars. Martian geography is mainly focused on what is called physical geography on Earth; that is the distribution of physical features across Mars and their cartographic representations.

History

Map of Mars by Giovanni Schiaparelli. North is at the top of this map; however, in most maps of Mars drawn before space exploration the convention among astronomers was to put south at the top because the telescopic image of a planet is inverted.

 

The first observations of Mars were from ground-based telescopes. The history of these observations are marked by the oppositions of Mars, when the planet is closest to Earth and hence is most easily visible, which occur every couple of years. Even more notable are the perihelic oppositions of Mars which occur approximately every 16 years, and are distinguished because Mars is closest to earth and Jupiter perihelion making it even closer to Earth.

In September 1877, (a perihelic opposition of Mars occurred on September 5), Italian astronomer Giovanni Schiaparelli published the first detailed map of Mars. These maps notably contained features he called canali ("channels"), that were later shown to be an optical illusion. These canali were supposedly long straight lines on the surface of Mars to which he gave names of famous rivers on Earth. His term was popularly mistranslated as canals, and so started the Martian canal controversy. 

Following these observations, it was a long-held belief that Mars contained vast seas and vegetation. It was not until spacecraft visited the planet during NASA's Mariner missions in the 1960s that these myths were dispelled. Some maps of Mars were made using the data from these missions, but it wasn't until the Mars Global Surveyor mission, launched in 1996 and ending in late 2006, that complete, extremely detailed maps were obtained. These maps are now available online at http://www.google.com/mars/

 

Topography

High resolution topographic map of Mars based on the Mars Global Surveyor laser altimeter research led by Maria Zuber and David Smith. North is at the top. Notable features include the Tharsis volcanoes in the west (including Olympus Mons), Valles Marineris to the east of Tharsis, and Hellas basin in the southern hemisphere.

 

STL 3D model of Mars with 20× elevation exaggeration using data from the Mars Global Surveyor Mars Orbiter Laser Altimeter.

 

Mars, 2001, with the southern polar ice cap visible on the bottom.

 

North Polar region with icecap.

 

Given that it is a planet, the geography of Mars varies considerably. However, the dichotomy of Martian topography is striking: northern plains flattened by lava flows contrast with the southern highlands, pitted and cratered by ancient impacts. The surface of Mars as seen from Earth is consequently divided into two kinds of areas, with differing albedo. The paler plains covered with dust and sand rich in reddish iron oxides were once thought of as Martian 'continents' and given names like Arabia Terra (land of Arabia) or Amazonis Planitia (Amazonian plain). The dark features were thought to be seas, hence their names Mare Erythraeum, Mare Sirenum and Aurorae Sinus. The largest dark feature seen from Earth is Syrtis Major Planum. 

The shield volcano, Olympus Mons (Mount Olympus), rises 22 km above the surrounding volcanic plains, and is the highest known mountain on any planet in the solar system. It is in a vast upland region called Tharsis, which contains several large volcanos. The Tharsis region of Mars also has the solar system's largest canyon system, Valles Marineris or the Mariner Valley, which is 4,000 km long and 7 km deep. Mars is also scarred by countless impact craters. The largest of these is the Hellas impact basin. See list of craters on Mars. 

Mars has two permanent polar ice caps, the northern one located at Planum Boreum and the southern one at Planum Australe. 

The difference between Mars' highest and lowest points is nearly 30 km (from the top of Olympus Mons at an altitude of 21.2 km to the bottom of the Hellas impact basin at an altitude of 8.2 km below the datum). In comparison, the difference between Earth's highest and lowest points (Mount Everest and the Mariana Trench) is only 19.7 km. Combined with the planets' different radii, this means Mars is nearly three times "rougher" than Earth.

The International Astronomical Union's Working Group for Planetary System Nomenclature is responsible for naming Martian surface features. 

Zero elevation

On Earth, the zero elevation datum is based on sea level (the geoid). Since Mars has no oceans and hence no 'sea level', it is convenient to define an arbitrary zero-elevation level or "vertical datum" for mapping the surface, called areoid.

The datum for Mars was defined initially in terms of a constant atmospheric pressure. From the Mariner 9 mission up until 2001, this was chosen as 610.5 Pa (6.105 mbar), on the basis that below this pressure liquid water can never be stable (i.e., the triple point of water is at this pressure). This value is only 0.6% of the pressure at sea level on Earth. Note that the choice of this value does not mean that liquid water does exist below this elevation, just that it could were the temperature to exceed 273.16 K (0.01 degrees C, 32.018 degrees F).

In 2001, Mars Orbiter Laser Altimeter data led to a new convention of zero elevation defined as the equipotential surface (gravitational plus rotational) whose average value at the equator is equal to the mean radius of the planet.

Zero meridian

Mars' equator is defined by its rotation, but the location of its prime meridian was specified, as was Earth's, by choice of an arbitrary point which was accepted by later observers. The German astronomers Wilhelm Beer and Johann Heinrich Mädler selected a small circular feature in the Sinus Meridiani ('Middle Bay' or 'Meridian Bay') as a reference point when they produced the first systematic chart of Mars features in 1830–32. In 1877, their choice was adopted as the prime meridian by the Italian astronomer Giovanni Schiaparelli when he began work on his notable maps of Mars. In 1909 the ephemeris makers decided that it was more important to maintain continuity of the ephemerides as a guide to observations and this definition was "virtually abandoned."

After the Mariner spacecraft provided extensive imagery of Mars, in 1972 the Mariner 9 Geodesy/Cartography Group proposed that the prime meridian passed through the center of a small 500 m diameter crater (named Airy-0), located in Sinus Meridiani along the meridian line of Beer and Mädler, thus defining 0.0° longitude with a precision of 0.001°. This model used the planetographic control point network developed by Merton Davies of the RAND Corporation.

As radiometric techniques increased the precision with which objects could be located on the surface of Mars, the center of a 500 m circular crater was considered to be insufficiently precise for exact measurements. The IAU Working Group on Cartographic Coordinates and Rotational Elements therefore recommended setting the longitude of the Viking 1 lander, for which there were extensive radiometric tracking data, as marking the standard longitude of 47.95137° west. This definition maintains the position of the center of Airy-0 at 0° longitude, within the tolerance of current cartographic uncertainties.

Martian dichotomy

Observers of Martian topography will notice a dichotomy between the northern and southern hemispheres. Most of the northern hemisphere is flat, with few impact craters, and lies below the conventional ‘zero elevation’ level. In contrast, the southern hemisphere is mountains and highlands, mostly well above zero elevation. The two hemispheres differ in elevation by 1 to 3 km. The border separating the two areas is very interesting to geologists. 

One distinctive feature is the fretted terrain. It contains mesas, knobs, and flat-floored valleys having walls about a mile high. Around many of the mesas and knobs are lobate debris aprons that have been shown to be rock-covered glaciers.

Other interesting features are the large river valleys and outflow channels that cut through the dichotomy.

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  Fresh Impact crater on Mars 3.7°N 53.4°E (November 19, 2013).
  
  
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  Fretted terrain of Ismenius Lacus showing flat floored valleys and cliffs. Photo taken with Mars Orbiter Camera (MOC) on the Mars Global Surveyor.
  
  
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  Enlargement of the photo on the left showing cliff. Photo taken with high resolution camera of Mars Global Surveyor (MGS).
  
  
  
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  View of lobate debris apron along a slope. Image located in Arcadia quadrangle. 
  
  
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  Place where a lobate debris apron begins. Note stripes which indicate movement. Image located in Ismenius Lacus quadrangle.
   
  

The northern lowlands comprise about one-third of the surface of Mars and are relatively flat, with occasional impact craters. The other two-thirds of the Martian surface are the southern highlands. The difference in elevation between the hemispheres is dramatic. Because of the density of impact craters, scientists believe the southern hemisphere to be far older than the northern plains. Much of heavily cratered southern highlands date back to the period of heavy bombardment, the Noachian.

Multiple hypotheses have been proposed to explain the differences. The three most commonly accepted are a single mega-impact, multiple impacts, and endogenic processes such as mantle convection. Both impact-related hypotheses involve processes that could have occurred before the end of the primordial bombardment, implying that the crustal dichotomy has its origins early in the history of Mars.

The giant impact hypothesis, originally proposed in the early 1980s, was met with skepticism due to the impact area's non-radial (elliptical) shape, where a circular pattern would be stronger support for impact by larger object(s). But a 2008 study provided additional research that supports a single giant impact. Using geologic data, researchers found support for the single impact of a large object hitting Mars at approximately a 45-degree angle. Additional evidence analyzing Martian rock chemistry for post-impact upwelling of mantle material would further support the giant impact theory.

Nomenclature

Early nomenclature

Although better remembered for mapping the Moon starting in 1830, Johann Heinrich Mädler and Wilhelm Beer were the first "areographers". They started off by establishing once and for all that most of the surface features were permanent, and pinned down Mars' rotation period. In 1840, Mädler combined ten years of observations and drew the first map of Mars ever made. Rather than giving names to the various markings they mapped, Beer and Mädler simply designated them with letters; Meridian Bay (Sinus Meridiani) was thus feature "a". 

Over the next twenty years or so, as instruments improved and the number of observers also increased, various Martian features acquired a hodge-podge of names. To give a couple of examples, Solis Lacus was known as the "Oculus" (the Eye), and Syrtis Major was usually known as the "Hourglass Sea" or the "Scorpion". In 1858, it was also dubbed the "Atlantic Canale" by the Jesuit astronomer Angelo Secchi. Secchi commented that it "seems to play the role of the Atlantic which, on Earth, separates the Old Continent from the New" —this was the first time the fateful canale, which in Italian can mean either "channel" or "canal", had been applied to Mars. 

In 1867, Richard Anthony Proctor drew up a map of Mars-based, somewhat crudely, on the Rev. William Rutter Dawes' earlier drawings of 1865, then the best ones available. Proctor explained his system of nomenclature by saying, "I have applied to the different features the names of those observers who have studied the physical peculiarities presented by Mars." Here are some of his names, paired with those later used by Schiaparelli in his Martian map created between 1877 and 1886. Schiaparelli's names were generally adopted and are the names actually used today.

Proctor nomenclature	Schiaparelli nomenclature
Kaiser Sea	Syrtis Major
Lockyer Land	Hellas Planitia
Main Sea	Lacus Moeris
Herschel II Strait	Sinus Sabaeus
Dawes Continent	Aeria and Arabia
De La Rue Ocean	Mare Erythraeum
Lockyer Sea	Solis Lacus
Dawes Sea	Tithonius Lacus
Madler Continent	Chryse Planitia, Ophir, Tharsis
Maraldi Sea	Maria Sirenum and Cimmerium
Secchi Continent	Memnonia
Hooke Sea	Mare Tyrrhenum
Cassini Land	Ausonia
Herschel I Continent	Zephyria, Aeolis, Aethiopis
Hind Land	Libya

Proctor's nomenclature has often been criticized, mainly because so many of his names honored English astronomers, but also because he used many names more than once. In particular, Dawes appeared no fewer than six times (Dawes Ocean, Dawes Continent, Dawes Sea, Dawes Strait, Dawes Isle, and Dawes Forked Bay). Even so, Proctor's names are not without charm, and for all their shortcomings they were a foundation on which later astronomers would improve.

Modern nomenclature

Planet Mars - Topographical Map (USGS; 2005)

 

Today, names of Martian features derive from a number of sources, but the names of the large features are derived primarily from the maps of Mars made in 1886 by the Italian astronomer Giovanni Schiaparelli. Schiaparelli named the larger features of Mars primarily using names from Greek mythology and to a lesser extent the Bible. Mars' large albedo features retain many of the older names, but are often updated to reflect new knowledge of the nature of the features. For example, 'Nix Olympica' (the snows of Olympus) has become Olympus Mons (Mount Olympus).

Large Martian craters are named after important scientists and science fiction writers; smaller ones are named after towns and villages on Earth.

Various landforms studied by the Mars Exploration Rovers are given temporary names or nicknames to identify them during exploration and investigation. However, it is hoped that the International Astronomical Union will make permanent the names of certain major features, such as the Columbia Hills, which were named after the seven astronauts who died in the Space Shuttle Columbia disaster.

The Case for Mars

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/The_Case_for_Mars

 

The Case for Mars: The Plan to Settle the Red Planet and Why We Must

Author	Robert Zubrin Richard Wagner Arthur C. Clarke
Language	English
Subject	Non-fiction Science
Publisher	Touchstone
Publication date	1996
Pages	368
ISBN	978-0684835501
OCLC	34906203
Dewey Decimal	919.9/2304-dc20
LC Class	QB641.Z83 1996

The Case for Mars: The Plan to Settle the Red Planet and Why We Must is a nonfiction science book by Robert Zubrin, first published in 1996, and revised and updated in 2011.

The book details Zubrin's Mars Direct plan to make the first human landing on Mars. The plan focuses on keeping costs down by making use of automated systems and available materials on Mars to manufacture the return journey's fuel in situ. The book also reveals possible Mars colony designs and weighs the prospects for a colony's material self-sufficiency and for the terraforming of Mars.

Mars Direct

The Mars Direct plan was originally detailed by Zubrin and David Baker in 1990. The Case for Mars is, according to Zubrin, a comprehensive condensation for laymen of many years' work and research. Chapters one and four deal with Mars Direct most completely.

Colonization

For Robert Zubrin, the attractiveness of Mars Direct does not rest on a single cost-effective mission. He envisions a series of regular Martian missions with the ultimate goal of colonization, which he details in the seventh through ninth chapters. As initial explorers leave hab-structures on the planet, subsequent missions become easier to undertake.

Large subsurface, pressurized habitats would be the first step toward human settlement; the book suggests they can be built as Roman-style atria underground with easily produced Martian brick. During and after this initial phase of habitat construction, hard-plastic radiation- and abrasion-resistant geodesic domes could be deployed on the surface for eventual habitation and crop growth. Nascent industry would begin using indigenous resources: the manufacture of plastics, ceramics and glass.

The larger work of terraforming requires an initial phase of global warming to release atmosphere from the regolith and to create a water cycle. Three methods of global warming are described in the work and, Zubrin suggests, are probably best deployed in tandem: orbital mirrors to heat the surface; factories on the surface to pump halocarbons into the atmosphere; and the seeding of bacteria which can metabolize water, nitrogen and carbon to produce ammonia and methane (these would aid in global warming). While the work of warming Mars is on-going, true colonization can begin.

The Case for Mars acknowledges that any Martian colony will be partially Earth-dependent for centuries. However, it suggests that Mars may be a profitable place for two reasons. First, it may contain concentrated supplies of metals of equal or greater value to silver which have not been subjected to millennia of human scavenging and may be sold on Earth for profit. Secondly, the concentration of deuterium – a possible fuel for commercial nuclear fusion – is five times greater on Mars. Humans emigrating to Mars thus have an assured industry and the planet will be a magnet for settlers as wage costs will be high. The book asserts that “the labor shortage that will prevail on Mars will drive Martian civilization toward both technological and social advances.” 

Wider considerations

While detailing the exploration and colonization, The Case for Mars also addresses a number of attendant scientific and political factors.

Risks confronted

The fifth chapter analyzes various risks that putatively rule out a long-term human presence on Mars. Zubrin dismisses the idea that radiation and zero-gravity are unduly hazardous. He claims that cancer rates do increase for astronauts who have spent extensive time in space, but only marginally. Similarly, while zero-gravity presents challenges, “near total recovery of musculature and immune system occurs after reentry and reconditioning to a one-gravity environment.” Furthermore, since his plan has the spacecraft spinning at the end of a long tether to create artificial gravity, worries about zero gravity do not apply to this mission in any case. Back-contamination – humans acquiring and spreading Martian viruses – is described as "just plain nuts", because there are no host organisms on Mars for disease organisms to have evolved.

In the same chapter, Zubrin decisively denounces and rejects suggestions that the Moon should be used as waypoint to Mars or as a training area. It is ultimately much easier to journey to Mars from low Earth orbit than from the Moon and using the latter as a staging point is a pointless diversion of resources. While the Moon may superficially appear a good place to perfect Mars exploration and habitation techniques, the two bodies are radically different. The Moon has no atmosphere, no analogous geology and a much greater temperature range and rotational period. Antarctica or desert areas of Earth provide much better training grounds at lesser cost. 

Viability

In the third and tenth chapters, The Case for Mars addresses the politics and costs of the ideas described. The authors argue that the colonization of Mars is a logical extension of the settlement of North America. They envision a frontier society, providing opportunities for innovation and social experimentation. 

Zubrin suggests three models to provide the will and capital to drive Mars exploration forward: the J.F.K. model, in which a far-sighted U.S. leader provides the funding and mobilizes national public opinion around the idea; the Sagan model, in which international co-operation is the driving force; and the Gingrich approach, which emphasizes incentives and even prizes for private sector actors who take on research and development tasks. In keeping with the third idea, Zubrin describes twelve challenges that address various aspects of the exploration program. A monetary prize – from five hundred million to twenty billion dollars – is offered to companies who successfully complete the challenges. 

The prize-based approach to hardware development has emerged within the private aeronautics community, though not yet on the scale envisioned by Zubrin. Ventures such as the Ansari X-Prize and Robert Bigelow's America's Space Prize seek low-cost spaceflight development through private enterprise, and crucially, for the attainment of very specific predetermined goals in order to win the prizes.

The underlying political and economic problems of raising sufficient capital for terraforming using halocarbon emissions is critiqued by John Hickman.

Translations

In 2017, a Russian translation of the book was published under the title of Курс на Марс (On Course for Mars) (ISBN 978-5-699-75295-9).

Rocket propellant (updated)

From Wikipedia, the free encyclopedia

https://en.wikipedia.org/wiki/Rocket_propellant

 

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 I_sp, 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 I_sp is at a premium and thrust to weight ratios are less relevant.

Dense propellant launch vehicles have a higher takeoff mass due to lower I_sp, 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.