112-509: Raptor 2: 141.1 Raptor 2: 1,630 kg (3,590 lb) Raptor is a family of rocket engines developed and manufactured by SpaceX . It is the third rocket engine in history designed with a full-flow staged combustion (FFSC) fuel cycle, and the first such engine to power a vehicle in flight. The engine is powered by cryogenic liquid methane and liquid oxygen , a mixture known as methalox . SpaceX's super-heavy-lift Starship uses Raptor engines in its Super Heavy booster and in
224-433: 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
336-409: 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"
448-656: A 200 (metric) tons engine aiming for roughly 300 bar chamber pressure. [...] If you had it at a high expansion ratio, has the potential to have a specific impulse of 380." SpaceX aimed at a lifetime of 1000 flights. In January 2016, the United States Air Force (USAF) awarded a US$ 33.6 million development contract to SpaceX to develop a Raptor prototype for use on the upper stage of the Falcon 9 and Falcon Heavy . The contract required double-matching funding by SpaceX of at least US$ 67.3 million . Engine testing
560-570: A Raptor Vacuum was on S25 during the second integrated flight test . Raptor 2 is a complete redesign of the Raptor 1 engine. The turbomachinery, chamber, nozzle, and electronics were all redesigned. Many flanges were converted to welds , while other parts were deleted. Simplifications continued after production began. On 10 February 2022, Musk showed Raptor 2 capabilities and design improvements. By 18 December 2021, Raptor 2 had started production. By November 2022, SpaceX produced more than one Raptor
672-406: A day and had created a stockpile for future launches. Raptor 2s are produced at SpaceX's McGregor engine development facility . Raptor 2s were achieving 230 t f (510,000 lbf ) of thrust consistently by February 2022. Musk indicated that production costs were approximately half that of Raptor 1. Raptor 3 is aimed to ultimately achieve 330 tf (3.2 MN ) of thrust in
784-417: A direct measure of the engine's effectiveness in converting propellant mass into forward momentum. The specific impulse in terms of propellant mass spent has units of distance per time, which is a notional velocity called the effective exhaust velocity . This is higher than the actual exhaust velocity because the mass of the combustion air is not being accounted for. Actual and effective exhaust velocity are
896-434: A given propellant, when paired with a given engine, can accelerate its own initial mass at 1 g. The longer it can accelerate its own mass, the more delta-V it delivers to the whole system. In other words, given a particular engine and a mass of a particular propellant, specific impulse measures for how long a time that engine can exert a continuous force (thrust) until fully burning that mass of propellant. A given mass of
1008-526: A heavier engine with a higher specific impulse may not be as effective in gaining altitude, distance, or velocity as a lighter engine with a lower specific impulse, especially if the latter engine possesses a higher thrust-to-weight ratio . This is a significant reason for most rocket designs having multiple stages. The first stage is optimised for high thrust to boost the later stages with higher specific impulse into higher altitudes where they can perform more efficiently. The most common unit for specific impulse
1120-470: 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
1232-439: 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
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#17327827527951344-548: 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. Specific impulse Specific impulse (usually abbreviated I sp )
1456-417: A more energy-dense propellant can burn for a longer duration than some less energy-dense propellant made to exert the same force while burning in an engine. Different engine designs burning the same propellant may not be equally efficient at directing their propellant's energy into effective thrust. For all vehicles, specific impulse (impulse per unit weight-on-Earth of propellant) in seconds can be defined by
1568-417: A much higher specific impulse than rocket engines. For air-breathing engines, only the fuel mass is counted, not the mass of air passing through the engine. Air resistance and the engine's inability to keep a high specific impulse at a fast burn rate are limiting factors to the propellant consumption rate. If it were not for air resistance and the reduction of propellant during flight, specific impulse would be
1680-421: A much larger specific impulse than a rocket; for example a turbofan jet engine may have a specific impulse of 6,000 seconds or more at sea level whereas a rocket would be between 200 and 400 seconds. An air-breathing engine is thus much more propellant efficient than a rocket engine, because the air serves as reaction mass and oxidizer for combustion which does not have to be carried as propellant, and
1792-424: 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
1904-506: A planned in-space raptor re-light was cancelled due to rolling during coast. In November 2016, Raptor was projected to power the proposed Interplanetary Transport System (ITS), in the early 2020s. Musk discussed two engines: a sea-level variant (expansion ratio 40:1) with thrust of 3,050 kN (690,000 lbf) at sea level for the first stage/booster, and a vacuum variant (expansion ratio 200:1) with thrust of 3,285 kN (738,000 lbf) in space. 42 sea-level engines were envisioned in
2016-400: A rocket can be defined in terms of thrust per unit mass flow of propellant. This is an equally valid (and in some ways somewhat simpler) way of defining the effectiveness of a rocket propellant. For a rocket, the specific impulse defined in this way is simply the effective exhaust velocity relative to the rocket, v e . "In actual rocket nozzles, the exhaust velocity is not really uniform over
2128-646: A second Raptor production facility, in central Texas near the existing rocket engine test facility . The facility would concentrate on serial production of Raptor 2, while the California facility would produce Raptor Vacuum and new/experimental Raptor designs. The new facility was expected to eventually produce 800 to 1000 rocket engines each year. In 2019 the (marginal) cost of the engine was stated to be approaching US$ 1 million . SpaceX planned to mass-produce up to 500 Raptor engines per year, each costing less than US$ 250,000 . Raptor has evolved significantly since it
2240-401: A smaller Raptor engine—with slightly over half as much thrust as the previous designs—would be used on the next-generation rocket, a 9 m (30 ft)-diameter launch vehicle termed Big Falcon Rocket (BFR) and later renamed Starship . The redesign was aimed at Earth-orbit and cislunar missions so that the new system might pay for itself , in part, through economic spaceflight activities in
2352-605: A thrust of 1 meganewton (220,000 lb f ) and used the SpaceX-designed SX500 alloy, created to contain hot oxygen gas in the engine at up to 12,000 pounds per square inch (830 bar; 83 MPa). It was tested on a ground test stand in McGregor , firing briefly. To eliminate flow separation problems while testing in Earth's atmosphere, the test nozzle expansion ratio was limited to 150. By September 2017,
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#17327827527952464-432: 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,
2576-484: Is a measure of how efficiently a reaction mass engine, such as a rocket using propellant or a jet engine using fuel, generates thrust . A propulsion system with a higher specific impulse uses the mass of the propellant more efficiently. In the case of a rocket, this means less propellant needed for a given delta- v , so that the vehicle attached to the engine can more efficiently gain altitude and velocity. For engines like cold gas thrusters whose reaction mass
2688-490: Is also ionized, which would interfere with radio communication with the rocket. Nuclear thermal rocket engines differ from conventional rocket engines in that energy is supplied to the propellants by an external nuclear heat source instead of the heat of combustion . The nuclear rocket typically operates by passing liquid hydrogen gas through an operating nuclear reactor. Testing in the 1960s yielded specific impulses of about 850 seconds (8,340 m/s), about twice that of
2800-450: Is also valid for air-breathing jet engines, but is rarely used in practice. (Note that different symbols are sometimes used; for example, c is also sometimes seen for exhaust velocity. While the symbol I sp {\displaystyle I_{\text{sp}}} might logically be used for specific impulse in units of (N·s )/(m·kg); to avoid confusion, it is desirable to reserve this for specific impulse measured in seconds.) It
2912-400: 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
3024-412: 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
3136-408: 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
3248-404: 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
3360-554: 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,
3472-423: 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
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3584-434: Is impractical. Lithium and fluorine are both extremely corrosive, lithium ignites on contact with air, fluorine ignites on contact with most fuels, and hydrogen, while not hypergolic, is an explosive hazard. Fluorine and the hydrogen fluoride (HF) in the exhaust are very toxic, which damages the environment, makes work around the launch pad difficult, and makes getting a launch license that much more difficult. The rocket exhaust
3696-497: Is inversely proportional to specific fuel consumption (SFC) by the relationship I sp = 1/( g o ·SFC) for SFC in kg/(N·s) and I sp = 3600/SFC for SFC in lb/(lbf·hr). An example of a specific impulse measured in time is 453 seconds, which is equivalent to an effective exhaust velocity of 4.440 km/s (14,570 ft/s), for the RS-25 engines when operating in a vacuum. An air-breathing jet engine typically has
3808-456: 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
3920-423: Is needed to produce a given thrust for a given time and the more efficient the propellant is. This should not be confused with the physics concept of energy efficiency , which can decrease as specific impulse increases, since propulsion systems that give high specific impulse require high energy to do so. Thrust and specific impulse should not be confused. Thrust is the force supplied by the engine and depends on
4032-406: 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
4144-421: Is only the fuel they carry, specific impulse is exactly proportional to the effective exhaust gas velocity. In an atmospheric context, specific impulse can include the contribution to impulse provided by the mass of external air that is accelerated by the engine, such as by fuel combustion or by external propeller. Jet engines and turbofans breathe external air for both combustion and bypass, and therefore have
4256-409: 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
4368-625: Is powered by subcooled liquid methane and subcooled liquid oxygen in a full-flow staged combustion (FFSC) cycle. FFSC is a twin-shaft staged combustion cycle that uses both oxidizer-rich and fuel-rich preburners. The cycle allows for the full flow of both propellants through the turbines without dumping any unburnt propellant overboard. FFSC is a departure from the more common "open-cycle" gas generator system and LOX/kerosene propellants used by its predecessor Merlin . Before 2014, no FFSC had ever been used in an actual flight and only two FFSC designs had progressed sufficiently to reach test stands:
4480-443: 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
4592-408: Is proportional to the effective exhaust velocity. A spacecraft without propulsion follows an orbit determined by its trajectory and any gravitational field. Deviations from the corresponding velocity pattern (these are called Δ v ) are achieved by sending exhaust mass in the direction opposite to that of the desired velocity change. When an engine is run within the atmosphere, the exhaust velocity
SpaceX Raptor - Misplaced Pages Continue
4704-410: Is reduced by atmospheric pressure, in turn reducing specific impulse. This is a reduction in the effective exhaust velocity, versus the actual exhaust velocity achieved in vacuum conditions. In the case of gas-generator cycle rocket engines, more than one exhaust gas stream is present as turbopump exhaust gas exits through a separate nozzle. Calculating the effective exhaust velocity requires averaging
4816-409: Is related to the thrust , or forward force on the rocket by the equation: F thrust = v e ⋅ m ˙ , {\displaystyle F_{\text{thrust}}=v_{\text{e}}\cdot {\dot {m}},} where m ˙ {\displaystyle {\dot {m}}} is the propellant mass flow rate, which is the rate of decrease of
4928-407: Is spread among the entire fuel flow, meaning that the preburner exhaust driving the propellant turbopumps is as cool as possible, even cooler than other closed engine cycles that only preburn one propellant. This contributes to a long engine life. In contrast, an open-cycle engine in which the preburner exhaust bypasses the main combustion chamber tries to minimize the amount of propellant fed through
5040-427: 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
5152-455: Is that it may be used for rockets, where all the reaction mass is carried on board, as well as airplanes, where most of the reaction mass is taken from the atmosphere. In addition, giving the result as a unit of time makes the result easily comparable between calculations in SI units, imperial units, US customary units or other unit framework. The English unit pound mass is more commonly used than
5264-443: 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
5376-401: Is the product of the average specific gravity of a given propellant mixture and the specific impulse. While less important than the specific impulse, it is an important measure in launch vehicle design, as a low specific impulse implies that bigger tanks will be required to store the propellant, which in turn will have a detrimental effect on the launch vehicle's mass ratio . Specific impulse
5488-481: Is the second, as values are identical regardless of whether the calculations are done in SI , imperial , or US customary units. Nearly all manufacturers quote their engine performance in seconds, and the unit is also useful for specifying aircraft engine performance. The use of metres per second to specify effective exhaust velocity is also reasonably common. The unit is intuitive when describing rocket engines, although
5600-423: 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
5712-414: Is used, impulse is divided by propellant weight (weight is a measure of force), resulting in units of time (seconds). These two formulations differ from each other by the standard gravitational acceleration ( g 0 ) at the surface of the earth. The rate of change of momentum of a rocket (including its propellant) per unit time is equal to the thrust. The higher the specific impulse, the less propellant
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#17327827527955824-402: 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
5936-497: The Falcon 9 and Falcon Heavy launch vehicles. Raptor was conceived to burn hydrogen and oxygen propellants as of 2009. SpaceX had a few staff working on the Raptor upper-stage engine at a low priority in 2011. In October 2012, SpaceX announced concept work on an engine that would be "several times as powerful as the Merlin 1 series of engines, and won't use Merlin's RP-1 fuel". In November 2012, Musk announced that SpaceX
6048-589: The Starship second stage . Starship missions include lifting payloads to Earth orbit and is also planned for missions to the Moon and Mars . The engines are being designed for reuse with little maintenance. Raptor is designed for extreme reliability, aiming to support airline-level of safety required by the point-to-point Earth transportation market. Gwynne Shotwell claimed that Raptor would be able to deliver "long life... and more benign turbine environments". Raptor
6160-404: 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
6272-402: 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
6384-404: The effective exhaust velocity while reducing the actual exhaust velocity. Again, this is because the mass of the air is not counted in the specific impulse calculation, thus attributing all of the thrust momentum to the mass of the fuel component of the exhaust, and omitting the reaction mass, inert gas, and effect of driven fans on overall engine efficiency from consideration. Essentially,
6496-408: The upper stage of Falcon 9 and Falcon Heavy . The first version was intended to operate at a chamber pressure of 250 bars (25 MPa; 3,600 psi). As of July 2022, chamber pressure had reached 300 bars in a test. In April 2024, Musk shared the performance achieved by SpaceX with the Raptor 1 engine (sea level 185 tf, Rvac 200 tf) and Raptor 2 engine (sea level 230 tf, Rvac 258 tf) along with
6608-603: 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
6720-613: The FTS was activated. On the third integrated flight test , all 33 booster engines once again remained lit until main engine cutoff (MECO), and then following hot-staging, 13 successfully relit to perform a boostback for full duration. On the booster's landing burn, only 3 engines of the planned 13 lit, with 2 shutting down rapidly, the other remained lit until a rapid unscheduled disassembly (RUD) occurred ~462 metres above sea level. The ship successfully kept all 6 engines lit until second stage / secondary engine cutoff (SECO) without issues, however
6832-760: The Soviet RD-270 project in the 1960s and the Aerojet Rocketdyne Integrated Powerhead Demonstrator in the mid-2000s. RS-25 engines (first used on the Space Shuttle ) used a simpler form of staged combustion cycle. Several Russian rocket engines, including the RD-180 and the RD-191 did as well. FFSC has the advantage that the energy produced by the preburners, and used to power the propellant pumps,
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#17327827527956944-693: The Space Shuttle engines. A variety of other rocket propulsion methods, such as ion thrusters , give much higher specific impulse but with much lower thrust; for example the Hall-effect thruster on the SMART-1 satellite has a specific impulse of 1,640 s (16.1 km/s) but a maximum thrust of only 68 mN (0.015 lbf). The variable specific impulse magnetoplasma rocket (VASIMR) engine currently in development will theoretically yield 20 to 300 km/s (66,000 to 984,000 ft/s), and
7056-408: The actual exhaust speed is much lower, so the kinetic energy the exhaust carries away is lower and thus the jet engine uses far less energy to generate thrust. While the actual exhaust velocity is lower for air-breathing engines, the effective exhaust velocity is very high for jet engines. This is because the effective exhaust velocity calculation assumes that the carried propellant is providing all
7168-419: The amount of reaction mass flowing through the engine. Specific impulse measures the impulse produced per unit of propellant and is proportional to the exhaust velocity. Thrust and specific impulse are related by the design and propellants of the engine in question, but this relationship is tenuous. For example, LH 2 /LO 2 bipropellant produces higher I sp but lower thrust than RP-1 / LO 2 due to
7280-465: 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
7392-407: 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
7504-465: The booster/sea-level configuration. As of August 2024, it had reached 280 tf. It weighs 1525 kg. Chamber pressure reached 350 bar . 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
7616-427: The burned fuel. Next, inert gases in the atmosphere absorb heat from combustion, and through the resulting expansion provide additional thrust. Lastly, for turbofans and other designs there is even more thrust created by pushing against intake air which never sees combustion directly. These all combine to allow a better match between the airspeed and the exhaust speed, which saves energy/propellant and enormously increases
7728-414: The combustion chamber, rather than Merlin's pintle injectors . Raptor is designed for deep cryogenic propellants—fluids cooled to near their freezing points , rather than their boiling points , as is typical for cryogenic rocket engines. Subcooled propellants are denser, increasing propellant mass as well as engine performance. Specific impulse is increased, and the risk of cavitation at inputs to
7840-424: 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,
7952-462: 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
8064-474: The definition of specific impulse as impulse per unit mass of propellant. Specific fuel consumption is inversely proportional to specific impulse and has units of g/(kN·s) or lb/(lbf·h). Specific fuel consumption is used extensively for describing the performance of air-breathing jet engines. Specific impulse, measured in seconds, can be thought of as how many seconds one kilogram of fuel can produce one kilogram of thrust. Or, more precisely, how many seconds
8176-703: The design evolved from the Mars Colonial Transporter to the Interplanetary Transport System , the Big Falcon Rocket , and ultimately, Starship. The concept evolved from a family of Raptor-designated rocket engines (2012) to focus on the full-size Raptor engine (2014). In January 2016, the US Air Force awarded a US$ 33.6 million development contract to SpaceX to develop a prototype Raptor for use on
8288-452: 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
8400-492: The effective exhaust speed of the engines may be significantly different from the actual exhaust speed, especially in gas-generator cycle engines. For airbreathing jet engines , the effective exhaust velocity is not physically meaningful, although it can be used for comparison purposes. Metres per second are numerically equivalent to newton-seconds per kg (N·s/kg), and SI measurements of specific impulse can be written in terms of either units interchangeably. This unit highlights
8512-638: The entire exit cross section and such velocity profiles are difficult to measure accurately. A uniform axial velocity, v e , is assumed for all calculations which employ one-dimensional problem descriptions. This effective exhaust velocity represents an average or mass equivalent velocity at which propellant is being ejected from the rocket vehicle." The two definitions of specific impulse are proportional to one another, and related to each other by: v e = g 0 ⋅ I sp , {\displaystyle v_{\text{e}}=g_{0}\cdot I_{\text{sp}},} where This equation
8624-435: The exhaust gases having a lower density and higher velocity ( H 2 O vs CO 2 and H 2 O). In many cases, propulsion systems with very high specific impulse—some ion thrusters reach 10,000 seconds—produce low thrust. When calculating specific impulse, only propellant carried with the vehicle before use is counted. For a chemical rocket, the propellant mass therefore would include both fuel and oxidizer . In rocketry,
8736-455: The following equation: F thrust = g 0 ⋅ I sp ⋅ m ˙ , {\displaystyle F_{\text{thrust}}=g_{0}\cdot I_{\text{sp}}\cdot {\dot {m}},} where: I sp in seconds is the amount of time a rocket engine can generate thrust, given a quantity of propellant whose weight is equal to the engine's thrust. The advantage of this formulation
8848-447: The high-level design of the first stage. Three gimbaled sea-level Raptor engines would be used for landing the second stage. Six additional, non-gimbaled, vacuum-optimized Raptors (Raptor Vacuum) would provide primary thrust for the second stage, for a total of nine engines. Raptor Vacuums were envisioned to contribute a specific impulse of 382 s (3,750 m/s), using a much larger nozzle . In September 2017 Musk said that
8960-411: 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
9072-536: 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
9184-490: The momentum of engine exhaust includes a lot more than just fuel, but specific impulse calculation ignores everything but the fuel. Even though the effective exhaust velocity for an air-breathing engine seems nonsensical in the context of actual exhaust velocity, this is still useful for comparing absolute fuel efficiency of different engines. A related measure, the density specific impulse , sometimes also referred to as Density Impulse and usually abbreviated as I s d
9296-531: The near-Earth space zone. With the much smaller launch vehicle, fewer Raptor engines would be needed. BFR was then slated to have 31 Raptors on the first stage and 6 on the second stage. By mid-2018, SpaceX was publicly stating that the sea-level Raptor was expected to have 1,700 kN (380,000 lbf) thrust at sea level with a specific impulse of 330 s (3,200 m/s), with a nozzle exit diameter of 1.3 m (4.3 ft). Raptor Vacuum would have specific impulse of 356 s (3,490 m/s) in vacuum and
9408-417: 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
9520-495: 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
9632-403: 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
9744-465: The only reaction mass is the propellant, so the specific impulse is calculated using an alternative method, giving results with units of seconds. Specific impulse is defined as the thrust integrated over time per unit weight -on-Earth of the propellant: I sp = v e g 0 , {\displaystyle I_{\text{sp}}={\frac {v_{\text{e}}}{g_{0}}},} where In rockets, due to atmospheric effects,
9856-431: 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
9968-449: The preburner, which is achieved by operating the turbine at its maximum survivable temperature. An oxygen-rich turbine powers an oxygen turbopump, and a fuel-rich turbine powers a methane turbopump. Both oxidizer and fuel streams are converted completely to the gas phase before they enter the combustion chamber . This speeds up mixing and combustion, reducing the size and mass of the required combustion chamber. Torch igniters are used in
10080-463: The preburners. Raptor 2's main combustion chamber uses an undisclosed ignition method that is allegedly less complex, lighter, cheaper, and more reliable than Merlin's. Engine ignition in Raptor Vacuum is handled by dual-redundant spark-plug lit torch igniters, which eliminate the need for Merlin's dedicated, consumable igniter fluid. Raptor 2 uses coaxial swirl injectors to admit propellants to
10192-407: 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
10304-409: 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
10416-414: The reaction mass and all the thrust. Hence effective exhaust velocity is not physically meaningful for air-breathing engines; nevertheless, it is useful for comparison with other types of engines. The highest specific impulse for a chemical propellant ever test-fired in a rocket engine was 542 seconds (5.32 km/s) with a tripropellant of lithium , fluorine , and hydrogen . However, this combination
10528-425: 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
10640-452: The rocket lifted off from the launch mount. The flight test was terminated after climbing to an altitude of ~39 km over the Gulf of Mexico. Multiple engines were out before the flight termination system (FTS) destroyed the booster and ship. On the second integrated flight test all 33 booster engines remained lit until boostback burn startup, and all six Starship engines remained lit until
10752-422: The same in rocket engines operating in a vacuum. The amount of propellant can be measured either in units of mass or weight. If mass is used, specific impulse is an impulse per unit of mass, which dimensional analysis shows to have units of speed, specifically the effective exhaust velocity . As the SI system is mass-based, this type of analysis is usually done in meters per second. If a force-based unit system
10864-520: 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
10976-574: The slug, and when using pounds per second for mass flow rate, it is more convenient to express standard gravity as 1 pound-force per pound-mass. Note that this is equivalent to 32.17405 ft/s2, but expressed in more convenient units. This gives: F thrust = I sp ⋅ m ˙ ⋅ ( 1 l b f l b m ) . {\displaystyle F_{\text{thrust}}=I_{\text{sp}}\cdot {\dot {m}}\cdot \left(1\mathrm {\frac {lbf}{lbm}} \right).} In rocketry,
11088-490: The specific impulse varies with altitude, reaching a maximum in a vacuum. This is because the exhaust velocity isn't simply a function of the chamber pressure, but is a function of the difference between the interior and exterior of the combustion chamber . Values are usually given for operation at sea level ("sl") or in a vacuum ("vac"). Because of the geocentric factor of g 0 in the equation for specific impulse, many prefer an alternative definition. The specific impulse of
11200-440: The speed of development and testing. The 2016 subscale development engine had 40% (by mass) of its parts manufactured by 3D printing. In 2019, engine manifolds were cast from SpaceX's in-house developed SX300 Inconel superalloy, later changed to SX500. SpaceX's Merlin and Kestrel rocket engines use a RP-1 and liquid oxygen ("kerolox") combination. Raptor has about triple the thrust of SpaceX's Merlin 1D engine, which powers
11312-412: 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
11424-500: The subscale engine had completed 1200 seconds of firings across 42 tests. SpaceX completed many static fire tests on a vehicle using Raptor 2s, including a 31 engine test (intended to be 33) on 9 February 2023, and a 33 engine test on 25 August 2023. During testing, more than 50 chambers melted, and more than 20 engines exploded. SpaceX completed its first integrated flight test of Starship on 20 April 2023. The rocket had 33 Raptor 2 engines, but three of those were shut down before
11536-454: The target specifications for the upcoming Raptor 3 (sea level 280 tf, Rvac 306 tf) and said SpaceX would aim to ultimately achieve over 330 tonnes of thrust on the sea-level booster engines. Raptor 1 and 2 engines require a heat shroud to protect pipes and wiring from the heat of high-velocity atmospheric re-entry, while Raptor 3 is designed so that it does not require an external heat shield. Initial development testing of Raptor components
11648-666: The turbopumps is reduced due to the higher propellant fuel mass flow rate per unit of power generated. Cavitation (bubbles) reduces fuel flow/pressure and can starve the engine, while eroding turbine blades. The oxidizer to fuel ratio of the engine is approximately 3.8 to 1. Methalox burns relatively cleanly, reducing carbon build-up in the engine. Liquid methane and oxygen propellants have been adopted by many companies, such as Blue Origin with its BE-4 engine, as well as Chinese startup Space Epoch's Longyun-70 . Many components of early Raptor prototypes were manufactured using 3D printing , including turbopumps and injectors, increasing
11760-415: The two mass flows as well as accounting for any atmospheric pressure. For air-breathing jet engines, particularly turbofans , the actual exhaust velocity and the effective exhaust velocity are different by orders of magnitude. This happens for several reasons. First, a good deal of additional momentum is obtained by using air as reaction mass, such that combustion products in the exhaust have more mass than
11872-446: The vehicle's mass. A rocket must carry all its propellant with it, so the mass of the unburned propellant must be accelerated along with the rocket itself. Minimizing the mass of propellant required to achieve a given change in velocity is crucial to building effective rockets. The Tsiolkovsky rocket equation shows that for a rocket with a given empty mass and a given amount of propellant, the total change in velocity it can accomplish
11984-565: Was done at NASA's Stennis Space Center , beginning in April 2014. Testing focused on startup and shutdown procedures, as well as hardware characterization and verification . SpaceX began testing injectors in 2014 and tested an oxygen preburner in 2015. 76 hot-fire tests of the preburner, totaling some 400 seconds of test time, were executed from April-August. By early 2016, SpaceX had constructed an engine test stand at their McGregor test site in central Texas for Raptor testing. The first Raptor
12096-519: Was expected to exert 1,900 kN (430,000 lbf) force with a specific impulse of 375 s (3,680 m/s), using a nozzle exit diameter of 2.4 m (7.9 ft). In the BFR update given in September 2018, Musk showed a video of a 71-second fire test of a Raptor engine, and stated that "this is Raptor that will power BFR, both the ship and the booster; it's the same engine. [...] approximately
12208-552: Was manufactured at the SpaceX Hawthorne facility in California. By August 2016 it was shipped to McGregor for development testing. The engine had 1 MN (220,000 lb f ) thrust. It was the first-ever FFSC methalox engine to reach a test stand. A subscale development engine was used for design validation. It was one-third the size of the engine designs that were envisioned for flight vehicles. It featured 200 bars (20 MPa; 2,900 psi) of chamber pressure, with
12320-618: Was planned for NASA's Stennis Space Center in Mississippi under US Air Force supervision. The USAF contract called for a single prototype engine and ground tests. In October 2017 USAF awarded a US$ 40.8 million modification contract for a Raptor prototype for the Evolved Expendable Launch Vehicle program. It was to use liquid methane and liquid oxygen propellants, a full-flow staged combustion cycle , and be reusable. In July 2021, SpaceX announced
12432-420: Was revealed. (kg) (t) Pressure (bar) Impulse (s) Engine Only Raptor Vacuum (RVac) is a variant of Raptor with an extended, regeneratively-cooled nozzle for higher specific impulse in space. The vacuum-optimized Raptor targets a specific impulse of ~380 s (3,700 m/s). A full-duration test of version 1 of Raptor Vacuum was completed in September 2020 at McGregor. The first in-flight ignition of
12544-650: Was working on methane -fueled rocket engines, that Raptor would be methane-based, and that methane would fuel Mars colonization. Because of the presence of underground water and carbon dioxide in Mars atmosphere , methane, a simple hydrocarbon , could be synthesized on Mars using the Sabatier reaction . NASA found in-situ resource production on Mars to be viable for oxygen, water, and methane production. In early 2014 SpaceX confirmed that Raptor would be used for both first and second stages of its next rocket. This held as
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