Rocket propellant performance

The prediction of rocket propellant specific impulse, as well as impulse under other conditions, may be reliably accomplished by calculation using as input the chemical composition, the heat of formation, and the density of the component propellant chemicals. Not only impulse but also the composition of exhaust species (and of species in the combustion chamber and the throat) may be calculated if the thermodynamic properties of the chemical species involved are known or can be estimated. The present standard computer code for such calculations is that described by Gordon and McBride.44 Theoretical performance predictions using such programs are widely used to guide propellant formulation efforts and to predict rocket propellant performance however, verification of actual performance is necessary. 

Experimental Determination of the Burning Rate. Experimental determinations of the burning rate are made with the closed tomb for gun propellants and the strand burner for rocket propellants. The closed bomb is essentially a heavy-wahed cylinder capable of withstanding pressures to 689 MPa (100,000 psi). It is equipped with a piezoelectric pressure gauge and the associated apparatus requited to measure the total chamber pressure, which is directly related to the force of the propellant. It also measures the rate of pressure rise as a function of pressure which can then be related to the linear burning rate of the propellant via its geometry. Other devices, such as the Dynagun and the Hi—Low bomb, have also been developed for the measurement of gun propellant performance. 

The combustion performance of a rocket motor is dependent on various physicochemical processes that occur during propellant burning. Since the free volume of a rocket motor is limited for practical reasons, the residence time of the reactive materials that produce the high temperature and high pressure for propulsion is too short to allow completion of the reaction within the limited volume of the motor as a reactor. Though rocket motor performance is increased by the addition of energetic materials such as nitramine particles or azide polymers, sufficient reaction time for the main oxidizer and fuel components is required. 

The specific impulse of a rocket motor, I, as defined in Eq. (1.75), is dependent on both propellant combushon efficiency and nozzle performance. Since is also defined by Eq. (1.79), rocket motor performance can also be evaluated in terms of the characterishc velocity, c, defined in Eq. (1.74) and the thrust coefficient, Cp, defined in Eq. (1.70). Since c is dependent on the physicochemical parameters in the combustion chamber, the combushon performance can be evaluated in terms of c. On the other hand, Cp is dependent mainly on the nozzle expansion process, and so the nozzle performance can be evaluated in terms of Cp. Experimental values of and Cpgxp are obtained from measurements of chamber pressure, p, and thrust, F ... 

Though the pyrolants used in gas-hybrid rockets burn in a similar manner as rocket propellants, their chemical compositions are fuel-rich. The pyrolants burn incompletely and the combustion temperature is below about 1000 K. When an atomized oxidizer is mixed with the fuel-rich gas in the secondary combustor, the mixture reacts to generate high-temperature combustion products. The combushon performance designated by specific impulse, is dependent on the combinahon of pyrolant and oxidizer. 

The performance of rocket propellants is commonly studied by means of the specific impulse which can be expressed as the thrust delivered per unit weight of propellant consumed as shown by equation (3.9). 

J.R. Muenger Sc L. Greiner, Estimation of Performance Factors for Rocket Propellants , Texaco Inc, Beacon, NY (1962) 293PP... 

P.I. Gold, Chemical species and chemical reactions of importance in nonequilibrium propellant performance calculations, NASA Accession No N66-33714, Rept No NASA-CR-65442, avail CFSTI, SciTechAerosp Rept 4 (19), 3722 (1966) CA 67, 4484 (1967) 47) S.S. Cherry L.J. van Nice, Pyrodynamics 6 (3-4), 275 (1969) CA 70, 98394 (1969) 48) R.E. Lo, Theoretical performance of the multicomponent rocket propellant system (oxygen, fluorine/ beryllium, lithium hydride)/hydrogen, Deutsche Versuchsanst Luft- und Raumfehri Rept11968, DLR-Mitt-68-21 (Ger), avail CFSTI, SciTech Aerosp Rept 7 (1), 161 (1969) CA 71,... 

The physical and chemical characteristics of these candidate liquid propellants vary widely. However, all of the liquids which have found application as rocket propellants have one common characteristic—they are designed to fit the particular requirements of at least one particular rocket engine and vehicle system. Obviously, few liquids initially fulfill the requirements of a propulsion system designed to perform a particular mission. Thus, various compromises must be undertaken between the... 

CL-20 was evaluated in both propellant and explosive formulations. A large number of CL-20-based PBXs are reported in the literature. A comparison of their VOD with that of the corresponding HMX-based formulations reveals a 12-15% higher energy potential of CL-20-based formulations. CL-20 is also a superior alternative to RDX and HMX for application in low-signature rocket propellants. CL-20-based propellants offer burn-rates much higher (-35-110%) than those of HMX-based propellants. The performance of CL-20-based propellants in terms of Isp is found to be higher than those of RDX-based propellants. 

The propellant is the most vital sub-system of a gun and rocket or missile system and accordingly, the performance of a gun, rocket or missile mainly depends on the propellant used. Some basic performance parameters are used to define the characteristics of these propellants and they are different for gun and rocket propellants. 

NANOCAT Superfine Iron Oxide(SFIO) is a novel bum-rate catalyst and performs superbly in solid rocket propellants based on ammonium perchlorate (AP). SFIO provides a higher burn rate and lower pressure exponent compared with commercial iron oxides at equal concentrations. Some characteristics of NANOCAT SFIO as a burn-rate catalyst are as follows ... 

The inhibited propellants after conditioning at ambient, cold (-40 °C) and hot (+60 °C) temperatures, should give successful performance during static evaluation. This is the most important test and even if a polymer meets all the requirements but fails in this test, it cannot be used for inhibition of rocket propellants [282]. 

The specific impulse Is is used to compare the performances of rocket propellants and is dependent on the thrust and flow rate of the gases through the nozzle as shown in Equation 8.4. 

Information on the performance of some solid rocket propellants is presented in Table 8.2. 

Liquid rocket propellants are subdivided into monopropellants and bipropellants. Monopropellants are liquids which burn in the absence of external oxygen. They have comparatively low energy and specific impulse and are used in small missiles which require low thrust. Hydrazine is currently the most widely used monopropellant however, hydrogen peroxide, ethylene oxide, isopropyl nitrate and nitromethane have all been considered or used as monopropellants. Information on the performance of some monopropellants is presented in Table 8.3. 

S.Gordon A.R.Glueck, "Theoretical Performance of Liquid Amnonia with Liquid Oxygen as a Rocket Propellant, NACA Report RME58A21(May 1958)(Conf)(Not used as a source of info) 86)P.H,Groggins, edit,... 

Specific impulse,, the universally accepted measure of rocket engine performance, can also be used to indicate the performance of propellants. The most commonly stated expansion ration is 1,000 14.7 giving sea-... 

Before any form of antimatter rocket can exist, a lightweight method must be developed for producing antiparticles at a flow rate of grains/second in contrast with the few dozen of antiparlicles produced in research laboratory generators. Also, a practical storage or containment method must arise inasmuch as antiparticles explode violently upon contact with normal matter. Reference 5 gives a performance estimate of an Ip of 3.06 x 10 seconds for a rocket propelled vehicle with a thrust/weight ratio of 10 1... 

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