Rocket motors combustion

Siloxanes, nevertheless, are used in small amts dissolved in N204 to reduce heat transfer thru rocket motor walls during combustion. The addn of 0.6 to 1.8% siloxane produces a heat-transfer reduction of greater than 30% (Ref 35a). Also, in the patent of Gordon et al (Ref 35b), rocket motor combustion stability is provided by the addn of a siloxane. A propint with a siloxane additive was found to bum with a nearly constant press of 600psig for approx 6 secs, whereas the same propint stabilized with cellulose acetate exhibits two press extremes and a continuously decreasing press after less... 

Since the rocket motor combustion process takes place at constant pressure with gases in equilibrium at the flame temperature, this above criterion is of the greatest interest here. 

Propellant Rocket Motor Combustion Instability, D. T. Harrje, Ed., Section 2.4, pp. 74-100, U.S. Govt. Printing OflRce, NASA SP-194, 1972. 

A steady-state rocket-type combustion spray unit has been developed, called high velocity oxy fuel (HVOF), that creates a steady state, continuous, supersonic spray stream (1.2—3 mm dia) resembling a rocket motor exhaust. The portable device injects and accelerates the particles inside a barrel (rocket nozzle). It produces coating quaHty and particle velocities equal to the D-gun at 5—10 times the spray rate with significantly reduced coating costs. 

A solid propellant is a mechanical (heterogeneous) or a chemical (homogeneous, or colloidal) mixture of solid-state fuel and oxidizer-rich chemicals. Specially-formed charges of solid propellant (grains) arc placed in the combustion chamber of the solid rocket motor (SRM) at a production facility. Once assembled, the engine does not require additional maintenance, making it simple, reliable and easy to use. 

Liner Formulations For. Polybutadiene Propellants Used In Small Rocket Applications , Report No DREU T.N. 1825/69, (Can) (1969) 7) A.K. Roberts, Preliminary Data On Unstable Combustion In Aluminized Polybutadiene Rocket Motors , Rept No DREU T.N. 1824/69 ... 

R.S. Brown et al, AdvanChemEng 7, 1—69 (1968) CA 72, 11368 (1970) The topics reviewed include types of solid proplnts, sohd-proplnt rocket motors, ignition, steady-state combustion, and combustion instability and termination... 

In a solid-propellant rocket motor, the propellant is contained within the wall of the combustion chamber, as shown in Fig. 1. This contrasts with liquid systems, where both the fuel and oxidizing components are stored in tanks external to the combustion chamber and are pumped or pressure-fed to the combustor. In hybrid systems, one component, usually the fuel, is contained in the combustion chamber, while the other component is fed to the chamber from a separate storage tank, as in liquid systems. The solid-propellant motor also has an ignition system located at one end to initiate operation of the rocket. The supersonic nozzle affects the conversion of... 

L. Crocco and S. Cheng. Theory of Combustion Instability in Liquid Propellant Rocket Motors. Butterworths, London, 1956. 

L. Crocco. Aspects of combustion instability in liquid propellant rocket motors, part 1. /. Am. Rocket Soc., 21 163-178, 1951. 

An ethylene oxide monopropellant rocket motor is considered part of a ram rocket power plant in which the turbulent exhaust of the rocket reacts with induced air in an afterburner. The exit area of the rocket motor is 8 cm2. After testing it is found that the afterburner length must be reduced by 42.3%. What size must the exit port of the new rocket be to accomplish this reduction with the same afterburner combustion efficiency The new rocket would operate at the same chamber pressure and area ratio. How many of the new rockets would be required to maintain the same level of thrust as the original power plant ... 

Since Cp indicates the efficiency of the expansion process in the nozzle flow and c indicates the efficiency of the combustion process in the chamber, gives an indication of the overall efficiency of a rocket motor. 

The chemical compositions and thermochemical properties of representative NC-NG and NC-TMETN double-base propellants are compared in Table 4.9. Though the NC/NG mass ratio of 0.80 is much smaller than the NC/TMETM mass ratio of 1.38, the combustion performance in terms of Tf and Mg is seen to be similar, and 0 is 109 kmol K kg for both propellants. In the case of rocket motor operation, Igp and pj, are also approximately equivalent for both propellants. 

Two types of igniters are used for ignition in solid rocket motors, those giving (i) high-volumetric and (ii) high-temperature combustion products. The combustion... 

The smoke characteristics of three types of pyrolants, namely nitropolymer pyrolants composed of NC-NG with and without a nickel catalyst, and a B-KNO3 pyrolant, have been examined in relation to the use of these pyrolants as igniters of rocket motors. Though nitropolymer pyrolants are fundamentally smokeless in nature, a large amount of black smoke is formed when they burn at low pressures below about 4 MPa due to incomplete combustion. Metallic nickel or organonickel compounds are known to catalyze the gas-phase reaction of nitropolymer pyrolants. 

When a composite propellant composed of ammonium perchlorate (AP) and a hydrocarbon polymer burns in a rocket motor, HCl, CO2, H2O, and N2 are the major combustion products and small amounts of radicals such as OH, H, and CH are also formed. These products are smokeless in nature and the formation of carbon particles is not seen. The exhaust plume emits weak visible light, but no afterburning occurs because AP composite propellants are stoichiometrically balanced mixtures and, in general, no diffusional flames are generated. 

Fig. 12.11 shows the structure of a rocket plume generated downstream of a rocket nozzle. The plume consists of a primary flame and a secondary flame.Fil The primary flame is generated by the exhaust combustion gas from the rocket motor without any effect of the ambient atmosphere. The primary flame is composed of oblique shock waves and expansion waves as a result of interaction with the ambient pressure. The structure is dependent on the expansion ratio of the nozzle, as described in Appendix C. Therefore, no diffusional mixing with ambient air occurs in the primary flame. The secondary flame is generated by mixing of the exhaust gas from the nozzle with the ambient air. The dimensions of the secondary flame are dependent not only on the combustion gas expelled from the exhaust nozzle, but also on the expansion ratio of the nozzle. A nitropolymer propellant composed of nc(0-466), ng(0-369), dep(0104), ec(0 029), and pbst(0.032) is used as a reference propellant to determine the effect of plume suppression. The burning rate characteristics of the propellants are shown in Fig. 6-31. Since the nitropolymer propellant is fuel-rich, the exhaust gas forms a combustible gaseous mixture with the ambient air. This gaseous mixture is ignited and afterburning occurs somewhat downstream of the nozzle exit. The major combustion products in the combustion chamber are CO, Hj, CO2, N2, and HjO. The fuel components are CO and H2, the mole fractions of which at the nozzle throat are co(0.47) and iH2(0.24).

Thus, AP is a valuable oxidizer for formulating smokeless propellants or smokeless gas generators. However, since the combustion products of AP composite propellants contain a relatively high concentration of hydrogen chloride (HCI), white smoke is generated when they are expelled from an exhaust nozzle into a humid atmosphere. When the HCI molecules diffuse into the air and collide with H2O molecules therein, an acid mist is formed which gives rise to visible white smoke. Typical examples are AP composite propellants used in rocket motors. Based on experimental observations, white smoke is formed when the relative humidity exceeds about 40 %. Thus, AP composite propellants without any metal particles are termed reduced-smoke propellants. On the other hand, a white smoke trail is always seen from the exhaust of a rocket projectile assisted by an aluminized AP composite propellant under any atmospheric conditions. Thus, aluminized AP composite propellants are termed smoke propellants. 

If the combustion products of a propellant attain a state of thermal equilibrium, the combustion temperature may be determined theoretically, as described in Chapter 2. However, the combustion in a rocket motor is incomplete and so the flame temperature remains below the adiabatic flame temperature.bl If one assumes that the flame temperature, T, varies with pressure, p, in a rocket motor, T is expressed byl5]... 

Combustion tests carried out for a rocket motor demonstrate a typical T combustion instability. Double-base propellants composed of NC-NG propellants with and without a catalyst (1 % nickel powder) were burned. Detailed chemical compositions of both propellants are given in Section 6.4.6 and the burning rate characteristics are shown in Fig. 6.29. The addition of nickel is seen to have no effect on burning rate and the pressure exponent is n = 0.70 for both propellants. 

The combustion tests conducted for a rocket motor show that the combustion becomes unstable below 1.7 MPa and that the burning acquires a chuffing mode in the case of the uncatalyzed propellant. However, as expected, the combustion is stable even below 0.5 MPa for the nickel-catalyzed NC-NG propellant, as shown in Fig. 13.13. Propellants for which the flame temperature decreases with decreasing pressure tend to exhibit T combustion instability. 

When an energetic material burns in a combustion chamber fitted with an exhaust nozzle for the combustion gas, oscillatory combustion occurs. The observed frequency of this oscillation varies widely from low frequencies below 10 Hz to high frequencies above 10 kHz. The frequency is dependent not only on the physical and chemical properties of the energetic material, but also on its size and shape. There have been numerous theoretical and experimental studies on the combustion instability of rocket motors. Experimental methods for measuring the nature of combustion instability have been developed and verified. However, the nature of combustion instability has not yet been fully understood because of the complex interactions between the combustion wave of propellant burning and the mode of acoustic waves. 

When combustion instability occurs for an internal burning grain of a rocket motor, the burning rate of the grain varies with time and so does the pressure in the rocket motor. The pressure versus time curve shows oscillations of a certain frequency. When the propellant burning mode is not in harmony with the pressure oscillation mode, the combustion instabiUty tends to decay. However, when the burning mode is in harmony with the oscillation mode, the pressure oscillation is amplified. 

Combustion of a propellant in a rocket motor accompanied by high-frequency pressure oscillation is one of the most harmful phenomena in rocket motor operation. There have been numerous theoretical and experimental studies on the acoustic mode of oscillation, concerning both the medium-frequency range of 100 Hz-1 kHz and the high-frequency range of 1 kHz-30 kHz. The nature of oscillatory combustion instability is dependent on various physicochemical parameters, such... 

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