Rocket motors, burning

Fig 17 Correlation of threshold pressure with propellant burning rate (5- x 40-inch rocket motor)... 

The discussion of the important design considerations of solid-propellant motors presented in Section I has shown the importance of the steady-state burning rate of the propellant. The particular mission for a rocket motor to... 

The rocket scientists wanted to be able to predict the thrust that could be expected from a fuel of a certain composition (see historical sketches by Zeleznik and Gordon, 1968 van Zeggeren and Storey, 1970 Smith and Missen, 1982). The volume of gases exiting the nozzle of the rocket motor could be used to calculate the expected thrust. The scientists recognized that by knowing the fuel s composition, the temperature at which it burned, and the pressure at the nozzle exit, they had uniquely defined the fuel s equilibrium volume, which they set about calculating. 

Heat and mass transfer through the boundary layer flow over the burning surface of propellants dominates the burning process for effechve rocket motor operation. Shock wave formahon at the inlet flow of ducted rockets is an important process for achieving high propulsion performance. Thus, a brief overview of the fundamentals of aerodynamics and heat transfer is provided in Appendices B -D as a prerequisite for the study of pyrodynamics. 

It was soon realized that platonized propellants, with their reduced temperature sensitivity in the plateau- and mesa-burning range, could be effectively used to minimize the sensitivity of the performance of a rocket to the temperature of the environment. Much work has been devoted to understanding the mechanism of plateau and mesa burning, with a view to optimizing the performance characteristics of rocket motors. 

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).

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. 

Fig. 13.21 shows another example of oscillatory burning of an RDX-AP composite propellant containing 0.40% A1 particles. The combustion pressure chosen for the burning was 4.5 MPa. The DC component trace indicates that the onset of the instability is 0.31 s after ignition, and that the instability lasts for 0.67 s. The pressure instability then suddenly ceases and the pressure returns to the designed pressure of 4.5 MPa. Close examination of the anomalous bandpass-filtered pressure traces reveals that the excited frequencies in the circular port are between 10 kHz and 30 kHz. The AC components below 10 kHz and above 30 kHz are not excited, as shown in Fig. 13.21. The frequency spectrum of the observed combustion instability is shown in Fig. 13.22. Here, the calculated frequency of the standing waves in the rocket motor is shown as a function of the inner diameter of the port and frequency. The sonic speed is assumed to be 1000 m s and I = 0.25 m. The most excited frequency is 25 kHz, followed by 18 kHz and 32 kHz. When the observed frequencies are compared with the calculated acoustic frequencies shown in Fig. 13.23, the dominant frequency is seen to be that of the first radial mode, with possible inclusion of the second and third tangential modes. The increased DC pressure between 0.31 s and 0.67 s is considered to be caused by a velocity-coupled oscillatory combustion. Such a velocity-coupled oscillation tends to induce erosive burning along the port surface. The maximum amplitude of the AC component pressure is 3.67 MPa between 20 kHz and 30 kHz. - ...

It is evident that the standing pressure wave in a rocket motor is suppressed by solid particles in the free volume of the combushon chamber. The effect of the pressure wave damping is dependent on the concentrahon of the solid parhcles, and the size of the parhcles is determined by the nature of the pressure wave, such as the frequency of the oscillation and the pressure level, as well as the properties of the combustion gases. Fig. 13.25 shows the results of combustion tests to determine the effechve mass fraction of A1 parhcles. When the propellant grain without A1 particles is burned, there is breakdown due to the combushon instability. When... 

End burning type rocket motor Fig. 14.1 Propellant burning in rocket motors. 

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