Liquid-Propellant Rocket Engine
Liquid-Propellant Rocket Engine
a reaction engine that runs on liquid propellant.
A diagram for a liquid-propellant rocket engine was developed in 1903 by K. E. Tsiolkovskii, who demonstrated the feasibility of using the engines in interplanetary flights. The principles of design proposed by Tsiolkovskii were supplemented by lu. V. Kondratiuk; the principles are still applied in the engines used today. The first liquid-propellant rocket engines were developed and tested by the American scientist R. Goddard in 1923 and by the German scientist H. Oberth in 1929. Work abroad on the development of the liquid-propellant rocket engine was conducted by the French scientist R. Esnault-Pelterie and by the German scientists E. Saenger and H. Walter. In the USSR the first liquid-propellant rocket engines were the ORM (“experimental rocket engine”) and the ORM-1, built and tested in the Gas Dynamics Laboratory in 1930–31 by V. P. Glushko. The OR-2 and the # 10 were developed by F. A. Tsander and the Group for the Study of Reactive Motion; they were tested in 1932–33.
A series of liquid-propellant rocket engines (ORM-1 to ORM-102) was built in the USSR during the 1930’s. These engines were used in testing the structural elements involved in ignition, start-up, and operation with a variety of liquid propellants, as well as in aircraft already in use (for example, the ORM-50 and the ORM-52).
In the 1940’s a great many types of liquid-propellant rocket engines were developed in the USSR and abroad both for a variety of rockets and for certain types of aircraft. The V-2 rocket (designed by W. von Braun) was first flight-tested in Germany in 1942 using a liquid-propellant engine, designed by W. Thiel, with a thrust of 245 kilonewtons (kN). Auxiliary liquid-propellant rocket engines, built by the Experimental Design Bureau (which had grown out of the Gas Dynamics Laboratory), were flight-tested between 1943 and 1946 in aircraft designed by V. M. Petliakov, S. A. Lavochkin, A. S. lakovlev, and P. O. Sukhoi.
Engines with much higher thrust were used in the USSR with ballistic missiles launched during the early 1950’s. Later, under the leadership of Glushko, A. M. Isaev, and S. A. Kosberg, Soviet designers developed and built engines that made possible the flights of the first Soviet-built artificial satellites of the earth, sun, moon, and Mars; of automatic stations on the moon, Venus, and Mars; of manned spacecraft; and of all geophysical and other rockets between 1949 and 1972. Liquid-propellant rocket engines have been developed widely in the United States, Great Britain, France, and elsewhere.
A liquid-propellant rocket engine consists of a combustion chamber with a nozzle, a feed system for the propellant components, control members, ignition devices, and auxiliary units (heat exchangers, mixers). Its thrust can range from millinewtons (in microrocket engines) to several meganewtons (the first-stage engine of the Saturn 5 generates a thrust of about 7 MN). The specific impulse may be as high as 4,500 N·sec/kg for bipropellants and 5,000 N·sec/kg for tripropellants. The engine’s mass-to-thrust ratio is between 0.7 and 2 g/N. The overall dimensions vary widely. The engines may be single-start or multiple-start and may have one or more chambers.
Rocket propulsion assemblies may have one or several engines. The propellant feed system of a liquid-propellant rocket engine may be of the displacement type, or it may be equipped with a turbopump unit (see Figure 1). Liquid-
propellant rocket engines with a turbopump have one of two basic operational schemes, either using afterburning of the generator gas or without afterburning. In liquid-propellant rocket engines equipped with a turbopump unit and without afterburning of the generator gas, the products of gas generation are first used in the turbine and are then discarded into the ambient through auxiliary nozzles (frequently used in steering). The generator gas, a product of incomplete combustion, has a relatively low temperature, so that the auxiliary nozzles exhibit a lesser degree of expansion than the main nozzles. Consequently, the specific impulse obtained by the discharge of the combustion products through the auxiliary nozzles is lower than the specific impulse of the main chamber of the engine; that is, there is a loss of specific impulse. In liquid-propellant rocket engines with afterburning of the generator gas, the products of gas generation obtained from the primary propellant components, having a relatively low temperature, are first used in the turbine and then pass into the engine chamber for afterburning. Engines of this type do not exhibit any loss of specific impulse attributable to the turbopump drive.
Liquid-propellant rocket engines can be categorized by purpose as (1) sustainer (main propulsion) engines, (2) vernier engines, (3) retrofire engines, and (4) control engines. Microrocket engines can be used for stabilization and orientation. Liquid-propellant rocket engines usually operate at a constant combustion-chamber pressure, although microrocket engines may be of the impulse type. Some hybrid engines now being developed, such as the turborocket and air-jet types, also use liquid-propellant rocket engines. Liquid-propellant rocket engines can also be classified, according to the type of oxidizer used, as nitric-acid, nitrogen-tetroxide, oxygen, hydrogen-peroxide, and fluorine engines.
Numerous problems arise in building liquid-propellant rocket engines. It is essential that the choice of propellant meet the required specific impulse and operating conditions and that operation achieve the necessary specific impulse. Steady operation must be achieved for the performance required, eliminating both low-frequency and high-frequency pressure oscillations, which can cause destructive engine vibrations. There are considerable difficulties in cooling an engine that has been exposed to the corrosive products of combustion at extremely high temperatures (to 5000°K) and pressures (to dozens of MN/m2); in some cases this is aggravated by the presence of a condensed phase. Most chambers are cooled by one of the propellant components; if this does not succeed in cooling the chamber and nozzle to the temperature necessary for reliability at full propellant usage, a lowered temperature is created in the gas layer adjacent to the wall by enriching the layer with one of the propellant components. Frequently, composite cooling is used (that is, internal and external cooling used simultaneously; see Figure 2). Thermal-protective coatings are widely used along with the cooling methods to protect the chamber walls and nozzles from heating. The reliability of the propellant feed system (cryogenic, corrosive)
is a complex problem at pressures in the dozens of MN/m2 and fuel consumption in the vicinity of several tons per second. Engine mass must be as low as possible.
REFERENCESTsiolkovskii, K. E. Issledovanie mirovykh prostranstv reaktivnymi priborami. Kaluga, 1926.
Dobrovol’skii, M. V.Zhidkostnye raketnye dvigateli. Moscow, 1968.
Alemasov, V. E., A. F. Dregalin, and A. P. Tishin. Teoriia raketnykh dvigatelei, 2nd ed. Moscow, 1969.
Petrovich, G. V. Raketnyie dvigateli GDL-OKB: 1929–1969. Moscow, 1969.
Volkov, E. B., L. G. Golovkov, and T. L. Syritsyn. Zhidkostnye raketnye dvigateli. Moscow, 1970.
Rocket Propulsion. Amsterdam-London-New York, 1960.
S. Z. KOPELEV