Rocket injector

Keywords Momentum ratio Pintle nozzle Rocket injector... 

Fig. 30.12 (a) Diameter and (b) velocity contours for two impinging jets of a rocket injector, where x is in the plane of the sheet and y is in the plane of impinging jets. The orifice diameters are 0.3 mm located 6.5 mm apart, making a 90 degrees angle with each other, with jet flow rate of 3.18ml/s [20] (Courtesy of American Institute of Physics)... 

Hautman, D.J., Spray Characterization of Like-on-Like Doublet Impinging Rocket Injectors, AIAA paper 91-0686, presented at 29th Aerospace Sciences Meeting, Rtmo, NV, Jan. 7—10, 1991. 

McHale, R.M., and Nurick, W.H., Noncircular Orifice Holes and Advanced Fabrication Techniques for Liquid Rocket Injectors Phase I Final Report, NASACR-108570, Rocketdyne Division of North American Rockwell Corp., Oct. 1970. 

Rocket Engine. A non-airbreathing reaction propulsion device that consists essentially of an injector, thrust chambers) and exhaust nozzle(s), and utilizes liquid fuels and oxidizers at controlled rates from which hot gases are generated by combustion and expanded thru a nozzle(s) (Ref 40a, p 125)... 

Thermal initiation of an explosion as well as the temperature of decomposition of propellants are measured by standard test No. 6 developed by the Joint Army-Navy-Air Force Panel on Liquid Propellants (2). The primary purpose of this test is to determine at what temperature an unstable material will undergo rapid exothermic decomposition. If the rate at which the heat is generated is greater than the rate at which it can be dissipated, an explosion will be likely. This test attempts to predict the result of conditions that can exist in the regenerative heating section of a rocket engine as well as at the propellant injector. Both of these sections experience relatively rapid local temperature rises owing to combustion. 

Injector design determines the physicochemical processes occurring in liquid propellant rocket engines. A complete quantitative description of the processes in liquid rockets is impossible because of our limited understanding of chemical reaction mechanisms and rates. The use of similarity principles simplifies the solution of theoretical combustion problems and is described for channel flow with chemical reactions and for diffusion flames over liquid droplets involving two coupled reaction steps. We find the new result that the observed burning rate of a liquid droplet is substantially independent of the relative rates of the coupled reactions. 

In conventional rocket engines, propellant distribution tends to be non-uniform across the injector face. Furthermore, relatively large changes in flow velocity may be associated with small fluctuations in supply pressures (4). Improved distribution of fuel and oxidizer across the injector face may be achieved by using orifice designs in which fully turbulent flow is attained reproducibly (4, 18). 

In summary, all features of the liquid rocket engine combustion processes are extensively affected by injector design, and any simplified combustion model, in which the essential three-dimensional nature of the flow processes is ignored, can only be of qualitative significance. Nevertheless, these simplified models are useful in giving us some insight into the nature of the physicochemical phenomena that determine engine performance. In this connection, steady-state combustion rates and overall combustion efficiencies in propellant utilization are far less important practical problems than are control or elimination of instabilities, excessive heat transfer, and hard starts. 

Burner Plates The term occasionally used by Brit rocket technologists for "injector ... 

The liquid rocket proplnts are introduced into the combustion chamber of a rocket through an injector in the form of droplets. The droplets then vaporize at their surface and the vapors start to burn. The mechanism of burning differs, however, depending upon whether the propint is a monopropellant or bipropellant, and, if a bipropellant, whether it is hypergolic or nonbypergolic... 

In rockets, the chamber pressure is the pressure developed as a result of burning of solid or liquid proplnts in the combustion chamber. It is usually measured thru the injector face or near the injector end of the thrust chamber by providing a small passage for the gas to press against some external measuring device(Ref 5)... 

Small-orifice injectors are used to atomize and mix the liquid propellants in appropriate proportions. The propellants enter the thrust chamber through the injection manifold and bum inside the thrust chamber. A typical liquid bipropellant rocket engine is shown in Fig. 37.20. 

Dickerson, R., Tate, K., and Barsic, N., Correlation of Spray Injector Parameters with Rocket Engine Performance, Technical Report AERPL-TR-68-147, Rocketd3Uie Division of North American Rockwell Corporation, Canoga Park, Calif., June, 1968. 

Abstract Pintle injectors have been developed for applications in rocket propulsion, but could potentially have high flowrate in other applications due to their relative simplicity. The spray from a pintle injector is formed from the collision of radial jets of fluid issuing from the center of the pintle post with an annular sleeve of fluid travelling axially along the post. The resultant interactions produce a conical spray similar to a hollow cone swirl atomizer. This chapter reviews historical applications of pintle injectors and theoretical bases for their design. 

The pintle injector is unique among the various injector options which have been successfully used in hquid rocket engines. Many early liquid engine injectors utilized impinging jets that create a spray due to the impact of the two streams, or coaxial injectors that mix propellants from shear or swirl induced in the inner fluid. These injector types are implemented in a flat face injector plate at the head-end of the combustion chamber. This resultant flowfield shown in Fig. 28.1c yields a curved spray/combustion zone which is substantially different than those formed by flat face-type injectors. 

Liquid Rocket Engine Injectors, NASA SP-8089, March 1976. 

EUco, E.R., and Stary, M.L., Final Report and Design Handbook of Acid-Aniline Rocket Motor and Injector Design, Report No. 455, Azusa (CA) Aerojet Engineering Corporation, June 9, 1950. 

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