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In rocket engine design, regenerative cooling is a configuration in which some or all of the propellant is passed through tubes, channels, or in a jacket around the combustion chamber or nozzle to cool the engine. This is effective because the propellants are often cryogenic. The heated propellant is then fed into a special gas-generator or injected directly into the main combustion chamber.
In 1857 Carl Wilhelm Siemens introduced the concept of regenerative cooling. [1] On 10 May 1898, James Dewar used regenerative cooling to become the first to statically liquefy hydrogen. [2] The concept of regenerative cooling was also mentioned in 1903 in an article by Konstantin Tsiolkovsky. [3] Robert Goddard built the first regeneratively cooled engine in 1923, but rejected the scheme as too complex. [4] A regeneratively cooled engine was built by the Italian researcher, Gaetano Arturo Crocco in 1930. The first Soviet engines to employ the technique were Fridrikh Tsander's OR-2 tested in March 1933 and the ORM-50, bench tested in November 1933 by Valentin Glushko. The first German engine of this type was also tested in March 1933 by Klaus Riedel in the VfR. The Austrian scientist Eugen Sänger was particularly famous for experiments with engine cooling starting in 1933; however, most of his experimental engines were water-cooled or cooled by an extra circuit of propellant.
The V-2 rocket engine, the most powerful of its time at 25 tons (245 kN) of thrust, was regeneratively cooled, in a design by Walter Thiel, by fuel pumped around the outside of the combustion chamber between the combustion chamber itself and an outer shell that conformed to the chamber and was separated by a few millimeters. This design was found to be insufficient to cool the combustion chamber due to the use of steel for the combustion chamber, and an additional system of fuel lines were added outside with connections through both combustion chamber shells to inject fuel directly into the chamber at an angle along the inner surface to further cool the chamber in a system called film cooling. This inefficient design required the burning of diluted alcohol at low chamber pressure to avoid melting the engine. The American Redstone engine used the same design.
A key innovation in regenerative cooling was the Soviet U-1250 engine designed by Aleksei Mihailovich Isaev in 1945. Its combustion chamber was lined with a thin copper sheet supported by the corrugated steel wall of the chamber. Fuel flowed through the corrugations and absorbed heat very efficiently. This permitted more energetic fuels and higher chamber pressures, and is the basic plan used in all Russian engines since. American engines usually solved this problem by lining the combustion chamber with brazed copper or nickel alloy tubes. Only recently engines like the RS-68 have started to use the cheaper Russian technique. The American style of lining the engine with copper tubes is called the "spaghetti construction", and the concept is credited to Edward A. Neu at Reaction Motors Inc. in 1947.
Regenerative cooling remains the predominant method for managing the thermal loads in thrust chambers. Typically the rocket fuel acts as a coolant as it enters the engine through passages at the nozzle exit. [5] It traverses the high-heat throat region and exits near the injector face. These passages are created either by brazing cooling tubes to the thrust chamber or by milling channels along the chamber walls. The cross-sections of these passages are smaller, increasing the coolant velocity and maximizing cooling efficiency in high-heat areas. [6]
The heat flux through the chamber wall is very high; usually in the range of 0.8–80 MW/m2 (0.5-50 BTU/in2-sec). [7] : 98 A common method for estimating the heat flux flowing out from the hot combustion gases is to use the Bartz equation: [8]
The amount of heat that can flow into the coolant is controlled by many factors including the temperature difference between the chamber and the coolant, the heat transfer coefficient, the thermal conductivity of the chamber wall, the velocity of the fluid inside the coolant channels, the velocity of the gas flow in the chamber/nozzle as well as the heat capacity and incoming temperature of the fluid used as a coolant.
Two boundary layers form: one in the hot gas in the chamber (which is modeled with the Bartz equation above) and the other in the coolant within the channels. [7] : 104–105
Very typically most of the temperature drop occurs in the gas boundary layer since gases are relatively poor conductors. This boundary layer can be destroyed however by combustion instabilities, and wall failure can follow very soon afterwards.
The boundary layer within the coolant channels can also be disrupted if the coolant is at subcritical pressure and film boils; the gas then forms an insulating layer and the wall temperature climbs very rapidly and soon fails. However, if the coolant engages in nucleate boiling but does not form a film, this helps disrupt the coolant boundary layer and the gas bubbles formed rapidly collapse; this can triple the maximum heat flow. However, many modern engines with turbopumps use supercritical coolants, and these techniques can be seldom used.
Regenerative cooling is seldom used in isolation; film cooling, [6] transpiration cooling, radiation cooling are frequently employed as well.
With regenerative cooling, the pressure in the cooling channels is greater than the chamber pressure. The inner liner is under compression, while the outer wall of the engine is under significant hoop stresses.
The metal of the inner liner is greatly weakened by the high temperature, and also undergoes significant thermal expansion at the inner surface while the cold-side wall of the liner constrains the expansion. This sets up significant thermal stresses that can cause the inner surface to crack or craze after multiple firings particularly at the throat.
In addition the thin inner liner requires mechanical support to withstand the compressive loading due to the propellant's pressure; this support is usually provided by the side walls of the cooling channels and the backing plate. The inner liner is usually constructed of relatively high temperature, high thermal conductivity materials; traditionally copper or nickel based alloys have been used.
Several different manufacturing techniques can be used to create the complex geometry necessary for regenerative cooling. These include a corrugated metal sheet brazed between the inner and outer liner; hundreds of pipes brazed into the correct shape, or an inner liner with milled cooling channels and an outer liner around that. [9] The geometry can also be created through direct metal 3D printing, as seen on some newer designs such as the SpaceX SuperDraco rocket engine.
A rocket is a vehicle that uses jet propulsion to accelerate without using any surrounding air. A rocket engine produces thrust by reaction to exhaust expelled at high speed. Rocket engines work entirely from propellant carried within the vehicle; therefore a rocket can fly in the vacuum of space. Rockets work more efficiently in a vacuum and incur a loss of thrust due to the opposing pressure of the atmosphere.
A heat exchanger is a system used to transfer heat between a source and a working fluid. Heat exchangers are used in both cooling and heating processes. The fluids may be separated by a solid wall to prevent mixing or they may be in direct contact. They are widely used in space heating, refrigeration, air conditioning, power stations, chemical plants, petrochemical plants, petroleum refineries, natural-gas processing, and sewage treatment. The classic example of a heat exchanger is found in an internal combustion engine in which a circulating fluid known as engine coolant flows through radiator coils and air flows past the coils, which cools the coolant and heats the incoming air. Another example is the heat sink, which is a passive heat exchanger that transfers the heat generated by an electronic or a mechanical device to a fluid medium, often air or a liquid coolant.
A scramjet is a variant of a ramjet airbreathing jet engine in which combustion takes place in supersonic airflow. As in ramjets, a scramjet relies on high vehicle speed to compress the incoming air forcefully before combustion, but where as a ramjet decelerates the air to subsonic velocities before combustion using shock cones, a scramjet has no shock cone and slows the airflow using shockwaves produced by its ignition source in place of a shock cone. This allows the scramjet to operate efficiently at extremely high speeds.
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 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 and rockets. Rocket vehicles carry their own oxidiser, unlike most combustion engines, so rocket engines can be used in a vacuum to propel spacecraft and ballistic missiles.
The expander cycle is a power cycle of a bipropellant rocket engine. In this cycle, the fuel is used to cool the engine's combustion chamber, picking up heat and changing phase. The now heated and gaseous fuel then powers the turbine that drives the engine's fuel and oxidizer pumps before being injected into the combustion chamber and burned.
RP-1 (alternatively, Rocket Propellant-1 or Refined Petroleum-1) is a highly refined form of kerosene outwardly similar to jet fuel, used as rocket fuel. RP-1 provides a lower specific impulse than liquid hydrogen (H2), but is cheaper, is stable at room temperature, and presents a lower explosion hazard. RP-1 is far denser than H2, giving it a higher energy density (though its specific energy is lower). RP-1 also has a fraction of the toxicity and carcinogenic hazards of hydrazine, another room-temperature liquid fuel.
A de Laval nozzle is a tube which is pinched in the middle, making a carefully balanced, asymmetric hourglass shape. It is used to accelerate a compressible fluid to supersonic speeds in the axial (thrust) direction, by converting the thermal energy of the flow into kinetic energy. De Laval nozzles are widely used in some types of steam turbines and rocket engine nozzles. It also sees use in supersonic jet engines.
A liquid-propellant rocket or liquid rocket utilizes a rocket engine burning liquid propellants. (Alternate approaches use gaseous or solid propellants.) Liquids are desirable propellants because they have reasonably high density and their combustion products have high specific impulse (Isp). This allows the volume of the propellant tanks to be relatively low.
The Aerojet Rocketdyne RS-25, also known as the Space Shuttle Main Engine (SSME), is a liquid-fuel cryogenic rocket engine that was used on NASA's Space Shuttle and is used on the Space Launch System (SLS).
The J-2, commonly known as Rocketdyne J-2, was a liquid-fuel cryogenic rocket engine used on NASA's Saturn IB and Saturn V launch vehicles. Built in the United States by Rocketdyne, the J-2 burned cryogenic liquid hydrogen (LH2) and liquid oxygen (LOX) propellants, with each engine producing 1,033.1 kN (232,250 lbf) of thrust in vacuum. The engine's preliminary design dates back to recommendations of the 1959 Silverstein Committee. Rocketdyne won approval to develop the J-2 in June 1960 and the first flight, AS-201, occurred on 26 February 1966. The J-2 underwent several minor upgrades over its operational history to improve the engine's performance, with two major upgrade programs, the de Laval nozzle-type J-2S and aerospike-type J-2T, which were cancelled after the conclusion of the Apollo program.
The pressure-fed engine is a class of rocket engine designs. A separate gas supply, usually helium, pressurizes the propellant tanks to force fuel and oxidizer to the combustion chamber. To maintain adequate flow, the tank pressures must exceed the combustion chamber pressure.
A rocket engine nozzle is a propelling nozzle used in a rocket engine to expand and accelerate combustion products to high supersonic velocities.
The YF-75 is a liquid cryogenic rocket engine burning liquid hydrogen and liquid oxygen in a gas generator cycle. It is China's second generation of cryogenic propellant engine, after the YF-73, which it replaced. It is used in a dual engine mount in the H-18 third stage of the Long March 3A, Long March 3B and Long March 3C launch vehicles. Within the mount, each engine can gimbal individually to enable thrust vectoring control. The engine also heats hydrogen and helium to pressurize the stage tanks and can control the mixture ratio to optimize propellant consumption.
A cold gas thruster is a type of rocket engine which uses the expansion of a pressurized gas to generate thrust. As opposed to traditional rocket engines, a cold gas thruster does not house any combustion and therefore has lower thrust and efficiency compared to conventional monopropellant and bipropellant rocket engines. Cold gas thrusters have been referred to as the "simplest manifestation of a rocket engine" because their design consists only of a fuel tank, a regulating valve, a propelling nozzle, and the little required plumbing. They are the cheapest, simplest, and most reliable propulsion systems available for orbital maintenance, maneuvering and attitude control.
Characteristic velocity or , or C-star is a measure of the combustion performance of a rocket engine independent of nozzle performance, and is used to compare different propellants and propulsion systems. c* should not be confused with c, which is the effective exhaust velocity related to the specific impulse by: . Specific impulse and effective exhaust velocity are dependent on the nozzle design unlike the characteristic velocity, explaining why C-star is an important value when comparing different propulsion system efficiencies. c* can be useful when comparing actual combustion performance to theoretical performance in order to determine how completely chemical energy release occurred. This is known as c*-efficiency.
A cryogenic rocket engine is a rocket engine that uses a cryogenic fuel and oxidizer; that is, both its fuel and oxidizer are gases which have been liquefied and are stored at very low temperatures. These highly efficient engines were first flown on the US Atlas-Centaur and were one of the main factors of NASA's success in reaching the Moon by the Saturn V rocket.
TM65 is a rocket engine developed by Copenhagen Suborbitals. TM65 uses Ethanol and liquid oxygen as propellants in a pressure-fed power cycle.
The TR-201 or TR201 is a hypergolic pressure-fed rocket engine used to propel the upper stage of the Delta rocket, referred to as Delta-P, from 1972 to 1988. The rocket engine uses Aerozine 50 as fuel, and N
2O
4 as oxidizer. It was developed in the early 1970s by TRW as a derivative of the lunar module descent engine (LMDE). This engine used a pintle injector first invented by Gerard W. Elverum Jr. and developed by TRW in the late 1950s and received US Patent in 1972. This injector technology and design is also used on SpaceX Merlin engines.
The RD-0110 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 supply vector control of the stage. It has an extensive flight history with its initial versions having flown more than 57 years ago.
The Bell Aerosystems Company XLR81 was an American liquid-propellant rocket engine, which was used on the Agena upper stage. It burned UDMH and RFNA fed by a turbopump in a fuel rich gas generator cycle. The turbopump had a single turbine with a gearbox to transmit power to the oxidizer and fuel pumps. The thrust chamber was all-aluminum, and regeneratively cooled by oxidizer flowing through gun-drilled passages in the combustion chamber and throat walls. The nozzle was a titanium radiatively cooled extension. The engine was mounted on a hydraulic actuated gimbal which enabled thrust vectoring to control pitch and yaw. Engine thrust and mixture ratio were controlled by cavitating flow venturis on the gas generator flow circuit. Engine start was achieved by solid propellant start cartridge.
