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75-704: The F-1 is a rocket engine developed by Rocketdyne . The engine uses a gas-generator cycle developed in the United States in the late 1950s and was used in the Saturn V rocket in the 1960s and early 1970s. Five F-1 engines were used in the S-IC first stage of each Saturn V, which served as the main launch vehicle of the Apollo program . The F-1 remains the most powerful single combustion chamber liquid-propellant rocket engine ever developed. Rocketdyne developed

150-433: 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

225-409: 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"

300-592: A competitor known as Pyrios , a liquid rocket booster , in NASA's Advanced Booster Program, which aims to find a more powerful successor to the five-segment Space Shuttle Solid Rocket Boosters intended for early versions of the Space Launch System. Pyrios uses two increased-thrust and heavily modified F-1B engines per booster. Due to the engine's potential advantage in specific impulse , if this F-1B configuration (using four F-1Bs in total) were integrated with

375-782: A ground-test replica, is on display as the first stage of a complete Saturn V at the Kennedy Space Center in Florida. SA-500D , the Dynamic Test Vehicle, is on display at the U.S. Space and Rocket Center in Huntsville, Alabama . A test engine is on display at the Powerhouse Museum in Sydney , Australia . It was the 25th out of 114 research and development engines built by Rocketdyne and it

450-541: A half minutes of operation, the five F-1s propelled the Saturn V vehicle to a height of 42 miles (222,000 ft; 68 km) and a speed of 6,164 mph (9,920 km/h). The combined flow rate of the five F-1s in the Saturn V was 3,357 US gal (12,710 L) or 28,415 lb (12,890 kg) per second. Each F-1 engine had more thrust than three Space Shuttle Main Engines combined. During static test firing,

525-470: 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

600-439: 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

675-880: A more northerly azimuth to reach a higher inclination orbit (50 degrees versus the usual 32.5 degrees). Ten F-1 engines were installed on two production Saturn Vs that never flew. The first stage from SA-514 is on display at the Johnson Space Center in Houston (although owned by the Smithsonian) and the first stage from SA-515 is on display at the INFINITY Science Center at John C. Stennis Space Center in Mississippi. Another ten engines were installed on two ground-test Saturn Vs never intended to fly. The S-IC-T "All Systems Test Stage,"

750-424: 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

825-512: A reduced number of engine parts, and the removal of the F-1 exhaust recycling system, including the turbine exhaust mid-nozzle and the "curtain" cooling manifold , with the turbine exhaust having a separate outlet passage beside the shortened main nozzle on the F-1B. The reduction in parts costs is aided by using selective laser melting in the production of some metallic parts. The resulting F-1B engine

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900-427: A second. The F-1 engine is the most powerful single-nozzle liquid-fueled rocket engine ever flown. The M-1 rocket engine was designed to have more thrust, but it was only tested at the component level. The later developed RD-170 is much more stable, technologically more advanced , more efficient and produces more thrust, but uses four nozzles fed by a single pump. The F-1 burned RP-1 (rocket grade kerosene ) as

975-432: 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,

1050-408: A way designed to promote mixing and combustion. Fuel was supplied to the injectors from a separate manifold; some of the fuel first traveled in 178 tubes down the length of the thrust chamber — which formed approximately the upper half of the exhaust nozzle — and back in order to cool the nozzle. A gas generator was used to drive a turbine which drove separate fuel and oxygen pumps, each feeding

1125-400: 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

1200-412: 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

1275-859: Is displayed vertically at the Museum of Flight in Seattle, Washington as part of the Apollo exhibit. An F-1 engine is installed vertically as a memorial to the Rocketdyne builders on De Soto Avenue, across the street from the former Rocketdyne plant in Canoga Park, California. It was installed in 1979, and moved from the parking lot across the street some time after 1980. An F-1 Engine is on display outside of The New Mexico Museum of Space History in Alamogordo, New Mexico. A recovered F-1 thrust chamber

1350-408: 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

1425-404: 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

1500-554: 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,

1575-423: 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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1650-633: Is intended to produce 1,800,000 lbf (8.0 MN) of thrust at sea level, a 15% increase over the approximate 1,550,000 lbf (6.9 MN) of thrust that the mature Apollo 15 F-1 engines produced. Sixty-five F-1 engines were launched aboard thirteen Saturn Vs, and each first stage landed in the Atlantic Ocean. Ten of these followed approximately the same flight azimuth of 72 degrees, but Apollo 15 and Apollo 17 followed significantly more southerly azimuths (80.088 degrees and 91.503 degrees, respectively). The Skylab launch vehicle flew at

1725-508: 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

1800-406: 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

1875-525: Is on display at the Cosmosphere . An intact engine (without nozzle extension) is displayed outdoors. On March 28, 2012, a team funded by Jeff Bezos , founder of Amazon.com , reported that they had located the F-1 rocket engines from an Apollo mission using sonar equipment. Bezos stated he planned to raise at least one of the engines, which rest at a depth of 14,000 feet (4,300 m), about 400 miles (640 km) east of Cape Canaveral, Florida. However,

1950-409: 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

2025-443: 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

2100-427: 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

2175-443: 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

2250-423: 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

2325-402: 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

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2400-651: The Marshall Space Flight Center began tests with an original F-1, serial number F-6049, which was removed from Apollo 11 due to a glitch. The engine was never used, and for many years it was at the Smithsonian Institution . The tests are designed to refamiliarize NASA with the design and propellants of the F-1 in anticipation of using an evolved version of the engine in future deep-space flight applications. In 2012, Pratt & Whitney , Rocketdyne , and Dynetics , Inc. presented

2475-600: The Museum of Flight in Seattle, WA and displays engine artifacts recovered including the thrust chamber and thrust chamber injector of the number 3 engine from the Apollo 12 mission, as well as a gas generator from an engine that powered the Apollo 16 flight. 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

2550-404: 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

2625-402: 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

2700-603: 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

2775-508: The F-1 and the E-1 to meet a 1955 U.S. Air Force requirement for a very large rocket engine. The E-1, although successfully tested in static firing, was quickly seen as a technological dead-end, and was abandoned for the larger, more powerful F-1. The Air Force eventually halted development of the F-1 because of a lack of requirement for such a large engine. However, the recently created National Aeronautics and Space Administration (NASA) appreciated

2850-558: The F-1, at 1,630,000 lbf (7.25 MN) per engine at sea level, however, each engine uses four combustion chambers instead of one, to solve the combustion instability problem. As part of the Space Launch System (SLS) program, NASA had been running the Advanced Booster Competition, which was scheduled to end with the selection of a winning booster configuration in 2015. In 2013, engineers at

2925-669: The Pyrios booster (see below) in 2013. As of 2013, none have proceeded beyond the initial study phase. The Comet HLLV would have used five F-1A engines on the main core and two on each of the boosters. The F-1 is the largest, highest-thrust single-chamber, single-nozzle liquid-fuel engine flown. Larger solid-fuel engines exist, such as the Space Shuttle Solid Rocket Booster with a sea-level liftoff thrust of 2,800,000 lbf (12.45 MN) apiece. The Soviet (now Russian) RD-170 can develop more thrust than

3000-486: The SLS Block 2, the vehicle could deliver 150 tonnes (330,000 lb) to low Earth orbit , while 130 tonnes (290,000 lb) is what is regarded as achievable with the planned solid boosters combined with a four-engine RS-25 core stage. The F-1B engine has a design goal to be at least as powerful as the unflown F-1A, while also being more cost effective. The design incorporates a greatly simplified combustion chamber,

3075-465: 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

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3150-407: 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

3225-424: 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,

3300-462: 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

3375-422: The condition of the engines, which had been submerged for more than 40 years, was unknown. NASA Administrator Charles Bolden released a statement congratulating Bezos and his team for their find and wished them success. He also affirmed NASA's position that any recovered artifacts would remain property of the agency, but that they would likely be offered to the Smithsonian Institution and other museums, depending on

3450-452: 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

3525-410: The engine was the thrust chamber, which mixed and burned the fuel and oxidizer to produce thrust. A domed chamber at the top of the engine served as a manifold supplying liquid oxygen to the injectors , and also served as a mount for the gimbal bearing which transmitted the thrust to the body of the rocket. Below this dome were the injectors, which directed fuel and oxidizer into the thrust chamber in

3600-451: The fuel and used liquid oxygen (LOX) as the oxidizer. A turbopump was used to inject fuel and oxygen into the combustion chamber. One notable challenge in the construction of the F-1 was regenerative cooling of the thrust chamber. Chemical engineer Dennis "Dan" Brevik was faced with the task of ensuring the preliminary combustion chamber tube bundle and manifold design produced by Al Bokstellar would run cool. In essence, Brevik's job

3675-490: The hot (5,800 °F (3,200 °C)) exhaust gas. Each second, a single F-1 burned 5,683 pounds (2,578 kg) of oxidizer and fuel: 3,945 lb (1,789 kg) of liquid oxygen and 1,738 lb (788 kg) of RP-1, generating 1,500,000 lbf (6.7 MN; 680 tf) of thrust. This equated to a flow rate of 671.4 US gal (2,542 L) per second; 413.5 US gal (1,565 L) of LOX and 257.9 US gal (976 L) of RP-1. During their two and

3750-411: 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

3825-536: 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

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3900-422: The kerosene-based RP-1 fuel left hydrocarbon deposits and vapors in the engine post test firing. These had to be removed from the engine to avoid problems during engine handling and future firing, and the solvent trichloroethylene (TCE) was used to clean the engine's fuel system immediately before and after each test firing. The cleaning procedure involved pumping TCE through the engine's fuel system and letting

3975-417: 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

4050-495: 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

4125-403: 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

4200-436: The number recovered. On March 20, 2013, Bezos announced he had succeeded in bringing parts of an F-1 engine to the surface, and released photographs. Bezos noted, "Many of the original serial numbers are missing or partially missing, which is going to make mission identification difficult. We might see more during restoration." The recovery ship was Seabed Worker , and had on board a team of specialists organized by Bezos for

4275-431: 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

4350-532: The post- Apollo era. However, the Saturn V production line was closed prior to the end of Project Apollo and no F-1A engines ever flew. There were proposals to use eight F-1 engines on the first stage of the Saturn C-8 and Nova rockets . Numerous proposals have been made from the 1970s and on to develop new expendable boosters based around the F-1 engine design. These include the Saturn-Shuttle , and

4425-407: 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

4500-535: The process of conservation. In August 2014, it was revealed that parts of two different F-1 engines were recovered, one from Apollo 11 and one from another Apollo flight, while a photograph of a cleaned-up engine was released. Bezos plans to put the engines on display at various places, including the National Air and Space Museum in Washington, D.C. On May 20, 2017, the Apollo permanent exhibit opened at

4575-409: 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

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4650-590: The recovery effort. On July 19, 2013, Bezos revealed that the serial number of one of the recovered engines is Rocketdyne serial number 2044 (equating to NASA number 6044), the #5 (center) engine that helped Neil Armstrong , Buzz Aldrin , and Michael Collins to reach the Moon with the Apollo 11 mission. The recovered parts were brought to the Kansas Cosmosphere and Space Center in Hutchinson for

4725-425: 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

4800-420: The running chamber responded to variations in pressure, and to determine how to nullify these oscillations. The designers could then quickly experiment with different co-axial fuel-injector designs to obtain the one most resistant to instability. These problems were addressed from 1959 through 1961. Eventually, engine combustion was so stable, it would self-damp artificially induced instability within one-tenth of

4875-520: 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

4950-485: The solvent overflow for a period ranging from several seconds to 30–35 minutes, depending upon the engine and the severity of the deposits. Sometimes the engine's gas generator and LOX dome were also flushed with TCE prior to test firing. The F-1 rocket engine had its LOX dome, gas generator, and thrust chamber fuel jacket flushed with TCE during launch preparations. Sources: F-1 thrust and efficiency were improved between Apollo 8 (SA-503) and Apollo 17 (SA-512), which

5025-412: 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

5100-487: The thrust chamber assembly. The turbine was driven at 5,500 RPM , producing 55,000 brake horsepower (41 MW). The fuel pump delivered 15,471 US gallons (58,560 litres) of RP-1 per minute while the oxidizer pump delivered 24,811 US gal (93,920 L) of liquid oxygen per minute. Environmentally, the turbopump was required to withstand temperatures ranging from input gas at 1,500 °F (820 °C) to liquid oxygen at −300 °F (−184 °C). Structurally, fuel

5175-681: The usefulness of an engine with so much power and contracted Rocketdyne to complete its development. Test firings of F-1 components had been performed as early as 1957. The first static firing of a full-stage developmental F-1 was performed in March 1959. The first F-1 was delivered to NASA MSFC in October 1963. In December 1964, the F-1 completed flight rating tests. Testing continued at least through 1965. Early development tests revealed serious combustion instability problems which sometimes caused catastrophic failure . Initially, progress on this problem

5250-512: Was 7,823,000 lbf (34.80 MN), which equates to an average F-1 thrust of 1,565,000 lbf (6.96 MN) – slightly more than the specified value. During the 1960s, Rocketdyne undertook uprating development of the F-1 resulting in the new engine specification F-1A. While outwardly very similar to the F-1, the F-1A produced about 20% greater thrust, 1,800,000 lbf (8 MN) in tests, and would have been used on future Saturn V vehicles in

5325-760: Was fired 35 times. The engine is on loan to the museum from the Smithsonian's National Air and Space Museum . It is the only F-1 on display outside the United States. An F-1 engine, on loan from the National Air and Space Museum, is on display at the Air Zoo in Portage, Michigan . An F-1 engine is on a horizontal display stand at Science Museum Oklahoma in Oklahoma City . F-1 engine F-6049

5400-420: Was necessary to meet the increasing payload capacity demands of later Apollo missions. There were small performance variations between engines on a given mission, and variations in average thrust between missions. For Apollo 15 , F-1 performance was: Measuring and making comparisons of rocket engine thrust is more complicated than it may first appear. Based on actual measurement the liftoff thrust of Apollo 15

5475-417: Was slow, as it was intermittent and unpredictable. Oscillations of 4 kHz with harmonics to 24 kHz were observed. Eventually, engineers developed a diagnostic technique of detonating small explosive charges (which they called "bombs") outside the combustion chamber, through a tangential tube ( RDX , C-4 or black powder were used) while the engine was firing. This allowed them to determine exactly how

5550-483: Was to "make sure it doesn’t melt." Through Brevik's calculations of the hydrodynamic and thermodynamic characteristics of the F-1, he and his team were able to fix an issue known as ‘starvation’. This is when an imbalance of static pressure leads to 'hot spots' in the manifolds. The material used for the F-1 thrust chamber tube bundle, reinforcing bands and manifold was Inconel-X750 , a refractory nickel based alloy capable of withstanding high temperatures. The heart of

5625-403: Was used to lubricate and cool the turbine bearings . Below the thrust chamber was the nozzle extension , roughly half the length of the engine. This extension increased the expansion ratio of the engine from 10:1 to 16:1. The exhaust from the turbine was fed into the nozzle extension by a large, tapered manifold; this relatively cool gas formed a film which protected the nozzle extension from

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