Rocket propellant is used as 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 .
94-426: The RD-0110 (or RO-8 , RD-0108 , RD-461 ) is a rocket engine burning liquid oxygen and kerosene in a gas generator combustion cycle . It has four fixed nozzles and the output of the gas generator is directed to four secondary vernier nozzles to provide attitude control for the stage. It has an extensive flight history with its initial versions having flown more than 64 years ago. OKB-154 of S.A. Kosberg
188-433: A propelling nozzle . The fluid is usually a gas created by high pressure (150-to-4,350-pound-per-square-inch (10 to 300 bar)) combustion of solid or liquid propellants , consisting of fuel and oxidiser components, within a combustion chamber . As the gases expand through the nozzle, they are accelerated to very high ( supersonic ) speed, and the reaction to this pushes the engine in the opposite direction. Combustion
282-409: A vacuum to propel spacecraft and ballistic missiles . Compared to other types of jet engine, rocket engines are the lightest and have the highest thrust, but are the least propellant-efficient (they have the lowest specific impulse ). The ideal exhaust is hydrogen , the lightest of all elements, but chemical rockets produce a mix of heavier species, reducing the exhaust velocity. Here, "rocket"
376-449: 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
470-408: 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. 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
564-470: A higher velocity compared to air. Expansion in the rocket nozzle then further multiplies the speed, typically between 1.5 and 2 times, giving a highly collimated hypersonic exhaust jet. The speed increase of a rocket nozzle is mostly determined by its area expansion ratio—the ratio of the area of the exit to the area of the throat, but detailed properties of the gas are also important. Larger ratio nozzles are more massive but are able to extract more heat from
658-439: A hot gas jet for propulsion. Alternatively, a chemically inert reaction mass can be heated by a high-energy power source through a heat exchanger in lieu of a combustion chamber. Solid rocket propellants are prepared in a mixture of fuel and oxidising components called grain , and the propellant storage casing effectively becomes the combustion chamber. Liquid-fuelled rockets force separate fuel and oxidiser components into
752-439: 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
846-622: A maximum limit determined only by the mechanical strength of the engine. In practice, the degree to which rockets can be throttled varies greatly, but most rockets can be throttled by a factor of 2 without great difficulty; the typical limitation is combustion stability, as for example, injectors need a minimum pressure to avoid triggering damaging oscillations (chugging or combustion instabilities); but injectors can be optimised and tested for wider ranges. Rocket propellant Rockets create thrust by expelling mass rear-ward, at high velocity. The thrust produced can be calculated by multiplying
940-424: A number called L ∗ {\displaystyle L^{*}} , the characteristic length : where: L* is typically in the range of 64–152 centimetres (25–60 in). The temperatures and pressures typically reached in a rocket combustion chamber in order to achieve practical thermal efficiency are extreme compared to a non-afterburning airbreathing jet engine . No atmospheric nitrogen
1034-571: 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
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#17327984965791128-698: 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. 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 and space applications. Their simplicity also makes solid rockets
1222-441: 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 ,
1316-415: 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
1410-445: 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 structural polymer. 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
1504-424: 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
1598-434: 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
1692-432: A variety of design approaches including turbopumps or, in simpler engines, via sufficient tank pressure to advance fluid flow. Tank pressure may be maintained by several means, including a high-pressure helium pressurization system common to many large rocket engines or, in some newer rocket systems, by a bleed-off of high-pressure gas from the engine cycle to autogenously pressurize the propellant tanks For example,
1786-412: 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
1880-432: 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
1974-400: Is designed for, but exhaust speeds as high as ten times the speed of sound in air at sea level are not uncommon. About half of the rocket engine's thrust comes from the unbalanced pressures inside the combustion chamber, and the rest comes from the pressures acting against the inside of the nozzle (see diagram). As the gas expands ( adiabatically ) the pressure against the nozzle's walls forces
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#17327984965792068-412: Is difficult to arrange in a lightweight fashion, although is routinely done with other forms of jet engines. In rocketry a lightweight compromise nozzle is generally used and some reduction in atmospheric performance occurs when used at other than the 'design altitude' or when throttled. To improve on this, various exotic nozzle designs such as the plug nozzle , stepped nozzles , the expanding nozzle and
2162-408: Is either measured as a speed (the effective exhaust velocity v e {\displaystyle v_{e}} in metres/second or ft/s) or as a time (seconds). For example, if an engine producing 100 pounds of thrust runs for 320 seconds and burns 100 pounds of propellant, then the specific impulse is 320 seconds. The higher the specific impulse, the less propellant is required to provide
2256-404: Is force divided by the rate of mass flow, this equation means that the specific impulse varies with altitude. Due to the specific impulse varying with pressure, a quantity that is easy to compare and calculate with is useful. Because rockets choke at the throat, and because the supersonic exhaust prevents external pressure influences travelling upstream, it turns out that the pressure at the exit
2350-554: Is ideally exactly proportional to the propellant flow m ˙ {\displaystyle {\dot {m}}} , provided the mixture ratios and combustion efficiencies are maintained. It is thus quite usual to rearrange the above equation slightly: and so define the vacuum Isp to be: where: And hence: Rockets can be throttled by controlling the propellant combustion rate m ˙ {\displaystyle {\dot {m}}} (usually measured in kg/s or lb/s). In liquid and hybrid rockets,
2444-423: Is important that the maximum pressures possible be created on the walls of the chamber and nozzle by a specific amount of propellant; as this is the source of the thrust. This can be achieved by all of: Since all of these things minimise the mass of the propellant used, and since pressure is proportional to the mass of propellant present to be accelerated as it pushes on the engine, and since from Newton's third law
2538-491: 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
2632-427: 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
2726-508: Is most frequently used for practical rockets, as the laws of thermodynamics (specifically Carnot's theorem ) dictate that high temperatures and pressures are desirable for the best thermal efficiency . Nuclear thermal rockets are capable of higher efficiencies, but currently have environmental problems which preclude their routine use in the Earth's atmosphere and cislunar space . For model rocketry , an available alternative to combustion
2820-406: Is no 'ram drag' to deduct from the gross thrust. Consequently, the net thrust of a rocket motor is equal to the gross thrust (apart from static back pressure). The m ˙ v e − o p t {\displaystyle {\dot {m}}\;v_{e-opt}\,} term represents the momentum thrust, which remains constant at a given throttle setting, whereas
2914-558: 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,
RD-0110 - Misplaced Pages Continue
3008-409: Is permitted to escape through an opening (the "throat"), and then through a diverging expansion section. When sufficient pressure is provided to the nozzle (about 2.5–3 times ambient pressure), the nozzle chokes and a supersonic jet is formed, dramatically accelerating the gas, converting most of the thermal energy into kinetic energy. Exhaust speeds vary, depending on the expansion ratio the nozzle
3102-443: Is present to dilute and cool the combustion, so the propellant mixture can reach true stoichiometric ratios. This, in combination with the high pressures, means that the rate of heat conduction through the walls is very high. In order for fuel and oxidiser to flow into the chamber, the pressure of the propellants entering the combustion chamber must exceed the pressure inside the combustion chamber itself. This may be accomplished by
3196-427: Is termed exhaust velocity , and after allowance is made for factors that can reduce it, the effective exhaust velocity is one of the most important parameters of a rocket engine (although weight, cost, ease of manufacture etc. are usually also very important). For aerodynamic reasons the flow goes sonic (" chokes ") at the narrowest part of the nozzle, the 'throat'. Since the speed of sound in gases increases with
3290-423: 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
3384-443: Is the water rocket pressurized by compressed air, carbon dioxide , nitrogen , or any other readily available, inert gas. Rocket propellant is mass that is stored, usually in some form of tank, or within the combustion chamber itself, prior to being ejected from a rocket engine in the form of a fluid jet to produce thrust. Chemical rocket propellants are the most commonly used. These undergo exothermic chemical reactions producing
3478-528: 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 (O 3 ), ClF 3 , and ClF 5 . 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. The theoretical exhaust velocity of
3572-416: 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,
3666-423: Is used as an abbreviation for "rocket engine". Thermal rockets use an inert propellant, heated by electricity ( electrothermal propulsion ) or a nuclear reactor ( nuclear thermal rocket ). Chemical rockets are powered by exothermic reduction-oxidation chemical reactions of the propellant: Rocket engines produce thrust by the expulsion of an exhaust fluid that has been accelerated to high speed through
3760-467: 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
3854-402: The A e ( p e − p a m b ) {\displaystyle A_{e}(p_{e}-p_{amb})\,} term represents the pressure thrust term. At full throttle, the net thrust of a rocket motor improves slightly with increasing altitude, because as atmospheric pressure decreases with altitude, the pressure thrust term increases. At the surface of
RD-0110 - Misplaced Pages Continue
3948-584: The Molniya-M Block-I, OKB-154 did an improvement program that put special emphasis on the reliability of the engine. This project gave birth to the RD-0110 . A particular problem observed during acceptance testing were high frequency instabilities, particularly during the start sequence. But it was solved by installing of longitudinal felt ribs on the combustion chamber. The development of the RD-0110
4042-404: The aerospike have been proposed, each providing some way to adapt to changing ambient air pressure and each allowing the gas to expand further against the nozzle, giving extra thrust at higher altitudes. When exhausting into a sufficiently low ambient pressure (vacuum) several issues arise. One is the sheer weight of the nozzle—beyond a certain point, for a particular vehicle, the extra weight of
4136-628: 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. 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)
4230-402: The combustion of reactive chemicals to supply the necessary energy, but non-combusting forms such as cold gas thrusters and nuclear thermal rockets also exist. Vehicles propelled by rocket engines are commonly used by ballistic missiles (they normally use solid fuel ) and rockets . Rocket vehicles carry their own oxidiser , unlike most combustion engines, so rocket engines can be used in
4324-419: 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
4418-409: 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,
4512-603: The Earth the pressure thrust may be reduced by up to 30%, depending on the engine design. This reduction drops roughly exponentially to zero with increasing altitude. Maximum efficiency for a rocket engine is achieved by maximising the momentum contribution of the equation without incurring penalties from over expanding the exhaust. This occurs when p e = p a m b {\displaystyle p_{e}=p_{amb}} . Since ambient pressure changes with altitude, most rocket engines spend very little time operating at peak efficiency. Since specific impulse
4606-455: 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. Although liquid hydrogen gives a high I sp , 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
4700-466: 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
4794-531: 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
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#17327984965794888-598: The Voronezh Mechanical Plant. Modifications to the RD-0107 design have led to production of four distinct versions of the engine: Rocket engine A rocket engine uses stored rocket propellants as the reaction mass for forming a high-speed propulsive jet of fluid, usually high-temperature gas. Rocket engines are reaction engines , producing thrust by ejecting mass rearward, in accordance with Newton's third law . Most rocket engines use
4982-465: The atmosphere, and while permitting the use of low pressure and hence lightweight tanks and structure. Rockets can be further optimised to even more extreme performance along one or more of these axes at the expense of the others. The most important metric for the efficiency of a rocket engine is impulse per unit of propellant , this is called specific impulse (usually written I s p {\displaystyle I_{sp}} ). This
5076-407: The axis of the engine, a side force may be imparted to the engine. This side force may change over time and result in control problems with the launch vehicle. Advanced altitude-compensating designs, such as the aerospike or plug nozzle , attempt to minimize performance losses by adjusting to varying expansion ratio caused by changing altitude. For a rocket engine to be propellant efficient, it
5170-432: 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
5264-424: The combustion chamber, where they mix and burn. Hybrid rocket engines use a combination of solid and liquid or gaseous propellants. Both liquid and hybrid rockets use injectors to introduce the propellant into the chamber. These are often an array of simple jets – holes through which the propellant escapes under pressure; but sometimes may be more complex spray nozzles. When two or more propellants are injected,
5358-462: The combustion gases, increasing the exhaust velocity. Vehicles typically require the overall thrust to change direction over the length of the burn. A number of different ways to achieve this have been flown: Rocket technology can combine very high thrust ( meganewtons ), very high exhaust speeds (around 10 times the speed of sound in air at sea level) and very high thrust/weight ratios (>100) simultaneously as well as being able to operate outside
5452-402: The corrugated jacket to save weight. All these sections use steel for construction. Given the extreme temperatures at the throat section, this part is made of copper alloy with milled channels and an external lining. A separate film cooling system is implemented though a different manifold and is injected though a circular slot upstream of the throat. The RD-0107/0108/0110 engines are produced in
5546-452: The desired impulse. The specific impulse that can be achieved is primarily a function of the propellant mix (and ultimately would limit the specific impulse), but practical limits on chamber pressures and the nozzle expansion ratios reduce the performance that can be achieved. Below is an approximate equation for calculating the net thrust of a rocket engine: Since, unlike a jet engine, a conventional rocket motor lacks an air intake, there
5640-414: The energy release per unit mass drops off quickly as the mixture ratio deviates from stoichiometric. LOX/LH 2 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/LH 2 rockets are generally limited in how rich they run by the performance penalty of
5734-400: The gas generator and the combustion chamber is done by pyrotechnic devices. The engine control is handled by a regulator, a throttle and a set of valves. It can throttle between 100% and 90.5%, with the option of 107% for a short time in emergencies. During development, combustion instability issues were observed. The problem was found to be intimately related to the injection system design. It
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#17327984965795828-494: 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: GOX (gaseous oxygen)
5922-445: 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
6016-438: The inner surface of the combustion chamber and this solved the issue permanently. The RD-0110 uses fuel as coolant for the regenerative cooling system. As most other Soviet designs, it uses a corrugated metal construction for the cooling jackets. The thrust chamber and upper nozzle sections has the corrugated metal sandwiched between an inner and outer metal layers. The lower section of the nozzle has no external lining, exposing
6110-411: The jet may be either below or above ambient, and equilibrium between the two is not reached at all altitudes (see diagram). For optimal performance, the pressure of the gas at the end of the nozzle should just equal the ambient pressure: if the exhaust's pressure is lower than the ambient pressure, then the vehicle will be slowed by the difference in pressure between the top of the engine and the exit; on
6204-536: The jets usually deliberately cause the propellants to collide as this breaks up the flow into smaller droplets that burn more easily. For chemical rockets the combustion chamber is typically cylindrical, and flame holders , used to hold a part of the combustion in a slower-flowing portion of the combustion chamber, are not needed. The dimensions of the cylinder are such that the propellant is able to combust thoroughly; different rocket propellants require different combustion chamber sizes for this to occur. This leads to
6298-605: The last Molniya flight was on 22 October 1967. For the crewed carrying Voskhod Block-I, a version of the engine that complied with the human rating 3K Regulations was needed. During 1963 OKB-154 developed this new version of the engine. Known with by the GRAU index 8D715P, the RD-0108 kept the same characteristics and performance of the RD-0107 while fully complying with the crew rating regulations. This engine had its first flight on 16 November 1963, and last flew on 29 June 1976. For
6392-578: The local rate of combustion. This positive feedback loop can easily lead to catastrophic failure of the case or nozzle. 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
6486-542: The main drivers of the design. For this reason a single shaft integrating the LOX pump, the RG-1 pump and the turbine was chosen. To provide acceptable suction performance, the RD-0110 turbopump has a dual inlet design with back to back centrifugal impellers. This allows it to work at relatively low inlet pressures without requiring additional booster pumps. The turbine is driven by a fuel rich gas generator. The ignition system for both
6580-430: 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: a storable oxidizer used with a solid fuel, which retains most virtues of both liquids (high ISP) and solids (simplicity). A hybrid-propellant rocket usually has
6674-440: 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:
6768-468: 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. 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
6862-417: The nozzle outweighs any performance gained. Secondly, as the exhaust gases adiabatically expand within the nozzle they cool, and eventually some of the chemicals can freeze, producing 'snow' within the jet. This causes instabilities in the jet and must be avoided. On a de Laval nozzle , exhaust gas flow detachment will occur in a grossly over-expanded nozzle. As the detachment point will not be uniform around
6956-495: The nozzle. As exit pressure varies from the ambient (atmospheric) pressure, a choked nozzle is said to be In practice, perfect expansion is only achievable with a variable–exit-area nozzle (since ambient pressure decreases as altitude increases), and is not possible above a certain altitude as ambient pressure approaches zero. If the nozzle is not perfectly expanded, then loss of efficiency occurs. Grossly over-expanded nozzles lose less efficiency, but can cause mechanical problems with
7050-403: The nozzle. Fixed-area nozzles become progressively more under-expanded as they gain altitude. Almost all de Laval nozzles will be momentarily grossly over-expanded during startup in an atmosphere. Nozzle efficiency is affected by operation in the atmosphere because atmospheric pressure changes with altitude; but due to the supersonic speeds of the gas exiting from a rocket engine, the pressure of
7144-431: The other hand, if the exhaust's pressure is higher, then exhaust pressure that could have been converted into thrust is not converted, and energy is wasted. To maintain this ideal of equality between the exhaust's exit pressure and the ambient pressure, the diameter of the nozzle would need to increase with altitude, giving the pressure a longer nozzle to act on (and reducing the exit pressure and temperature). This increase
7238-426: 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
7332-407: The pressure that acts on the engine also reciprocally acts on the propellant, it turns out that for any given engine, the speed that the propellant leaves the chamber is unaffected by the chamber pressure (although the thrust is proportional). However, speed is significantly affected by all three of the above factors and the exhaust speed is an excellent measure of the engine propellant efficiency. This
7426-409: The propellant flow entering the chamber is controlled using valves, in solid rockets it is controlled by changing the area of propellant that is burning and this can be designed into the propellant grain (and hence cannot be controlled in real-time). Rockets can usually be throttled down to an exit pressure of about one-third of ambient pressure (often limited by flow separation in nozzles) and up to
7520-446: 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
7614-509: 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
7708-412: 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
7802-425: The rocket engine in one direction while accelerating the gas in the other. The most commonly used nozzle is the de Laval nozzle , a fixed geometry nozzle with a high expansion-ratio. The large bell- or cone-shaped nozzle extension beyond the throat gives the rocket engine its characteristic shape. The exit static pressure of the exhaust jet depends on the chamber pressure and the ratio of exit to throat area of
7896-430: 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
7990-428: The rocket stage. Molecules with fewer atoms (like CO and H 2 ) have fewer available vibrational and rotational modes than molecules with more atoms (like CO 2 and H 2 O). 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
8084-520: The self-pressurization gas system of the SpaceX Starship is a critical part of SpaceX strategy to reduce launch vehicle fluids from five in their legacy Falcon 9 vehicle family to just two in Starship, eliminating not only the helium tank pressurant but all hypergolic propellants as well as nitrogen for cold-gas reaction-control thrusters . The hot gas produced in the combustion chamber
8178-414: 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 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
8272-412: The square root of temperature, the use of hot exhaust gas greatly improves performance. By comparison, at room temperature the speed of sound in air is about 340 m/s while the speed of sound in the hot gas of a rocket engine can be over 1700 m/s; much of this performance is due to the higher temperature, but additionally rocket propellants are chosen to be of low molecular mass, and this also gives
8366-402: 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
8460-478: 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
8554-460: Was finally solved by developing an optimized bi-propellant centrifugal atomizer design. During certification testing, high frequency combustion instabilities at start up were still observed. Even though the start instabilities were relatively rare at 1 in 60 to 80, and only on acceptance bench, great effort was made to eliminate the issue. Thanks to acoustic studies and modelling, a solution was found. Six combustible longitudinal felt ribs were placed at
8648-549: Was performed in 9 months during 1963, with its inaugural flight happening in 1964. It is also used on the Soyuz third stage in all models until the RD-0124 debut on the 2.1b . It has flown over 1350 times, and accumulated more than 336,500 s of burn time and is still flown many times per year. The RD-0110 was created in a period when KBKhA had just started to design rocket engines. After many studies, simplicity and reliability were
8742-662: Was tasked with developing an engine for the unmanned Molniya Block-I stage. Thus, the RD-0107 was developed in the 1960 to 1961 period, based on the RD-0106 ( GRAU Index: 8D715) engine that powered the SS-8 Sasin ICBM , also designed by OKB-154. It also leveraged the experience in the field from the Vostok Block-E RD-0105 / RD-0109 development. The engine had its debut flight on 10 October 1960, and
8836-570: Was used as the oxidizer for the Buran program 's orbital maneuvering system. 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 thrusters ionize a neutral gas and create thrust by accelerating the ions (or the plasma) by electric and/or magnetic fields. Thermal rockets use inert propellants of low molecular weight that are chemically compatible with
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