Rocket motors 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 ... 

Although analyses in which some of these assumptions are removed, for example, by accounting for finite gas-phase reaction rates [53], for non-uniform particle-size distributions [54], [55], [56] and for droplet breakup [57], [58], are more realistic in that correlations with observed rocket-motor performance sometimes can be obtained, they involve numerical integrations which may tend to obscure the essential ideas. Simple analytical results have, however, been 5eveloped for a model that accounts in an approximate way for nonuniform size distributions [59]. Comprehensive reviews of related studies may be found in [58] and [60]. The major drawback to all of these analyses is the one-dimensional flow approximation, which excludes from consideration the three-dimensional flows generally observed in real engines. 

Another key issue is the sensitivity of the mass flux or burning rate and surface temperature to the independent variables, P, To, qr. The sensitivity of burning rate to pressure and initial temperature is obviously important for internal ballistics and rocket motor performance prediction for quasi-static operation. The sensitivity of surface temperature is not quite as obvious but is related to the unsteady combustion behavior through Zeldovich-Novozhilov (ZN) theory. The sensitivity parameters are derivatives of the steady equations as defined in the nomenclature. Equation (15) or (16) can be differentiated with respect to initial temperature (To), pressure (Dg), and radiative flux (qr) to give... 

These fibres are used for their thermal properties combined with high mechanical performances. Unfortunately, their price is prohibitive and applications are reduced to, for example, rocket motors. 

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. 

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. 

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. 

Hercopel a unique all-epoxide cure composite solid propellant with excellent mechanical and ballistic properties. Its outstanding performance in extended environments makes it well suited for tactical missiles Double-Base Solid Propellants a wide variety of physical and ballistic properties which can be tailored to meet specific performance requirements. Their high specific impulse and excellent reproducibility are two of the many reasons Hercules double-base propellants are found in many of our rocket motors and gas generators used for both military and space applications... 

The standard ASTM D2585 filament wound pressurized bottle test method utilizes a 0.15-m (5.75-in.)intemal diameter filament wound bottle as the test article. This standard test method (with variation in bottle sizes) has been used extensively by the rocket motor industry [47-50] to evaluate glass, aramid, and graphite fiber composite vessel performance. This test method has generally shown good results, but is a relatively expensive test method. Testing of one 0.5-m (20-in.) diameter bottle can cost up to 20K. Other disadvantages are ... 

In a nutshell, ADN appears to be an efficient oxidizer for the high performance eco-friendly propellants and is now envisaged as a suitable and better successor to AP. The use of ADN in composite solid propellants eliminates the emission of chlorinated exhaust products from rocket motors and gives 5-10 s more Isp than conventional AP-based propellants. 

The parameters most commonly used to evaluate the performance of rocket engines are introduced first. From these, the significant parameters which determine the performance of propellants are derived. Similar to the approach in (15) simplified expressions will be derived theoretically in terms of the thermodynamic and other properties of the system in order to give insight into the fundamental significance of the individual performance parameters. These simplified expressions are derived from the so-called ideal rocket motor analysis. More accurate derivations are deferred until the next sections where the appropriate aspects of chemical thermodynamics are developed. 

The performance analysis of a rocket motor comprises the calculation of ... 

The h-S diagram becomes most convenient in following rocket motor processes and this is the reason for its introduction. The conveniences obtained are generally hidden by machine computation programs which essentially deal with the enthalpy-entropy process for the expansion process, h is the sensible enthalpy only. Theoretical performance calculations are performed in terms of the total enthalpy which is here defined as the sum of the sensible and chemical enthalpies only. 

Equilibrium concentrations of carbon or ammonia are not found in short combustion chambers used in rocket motors. The reason for this non-equilibrium situation is that the rate of formation of soot is very slow and carbon does not have time to form. Similarly the dissociation of NH3 is very slow. Thus in ethylene oxide monopropellant rocket motors one finds very little carbon, whereas equilibrium considerations predict carbon as a predominant product and in hydrazine decomposition chambers one finds an excess of NH3 over that predicted by equilibrium considerations. In ethylene oxide motors carbon forms from the decomposition of methane, not the reaction represented above, thus both non-equilibrium situations give higher performance than expected, since the endothermic reactions do not have time to take place. Of course, carbon also could form in cool reactions which take place in boundary layers along the walls where velocities are slow. 

Procedures for calculating the theoretical flame, or product temperature and product composition of a propellant mixture were discussed in the previous section. What remains to be analyzed is the nozzle expansion process. Since most thermochemical performance calculations are made in order to compare various propellants or propellant combinations, certain.ideal assumptions as discussed in Section n. A. are made. These ideal assumptions, however, only relate to the physical processes which actually occur in the rocket motor. As explained, normal dissipative losses such as friction and heat transfer are ignored. The gasses are assumed to enter the nozzle at zero velocity at the temperature and product composition calculated theoretically for the given mixture ratio of the propellant combination. 

A solid propellant rocket motor is quite simple in concept, although in practice a complete motor is more complex. As shown in Fig. 37.15, the rocket propellant is contained within a case, which may be metal or a reinforced high-performance composite. Frequently, the case is internally shielded by a bonded layer of insulation. The insulation is coated with a liner that bonds the propellant to the insulation. The integrity of the propellant-to-liner bond is of utmost importance failure at this interface during a motor firing can result in a sudden increase in the area of propellant surface exposed to combustion, with potentially catastrophic results. 

The length of time of storage during which an explosive material, generator, rocket motor, or component retains adequate performance characteristics under specified environmental conditions. 

The specific impulse of a propellant or a pair of reacting liquids in rocket motors is the most important characteristic of the performance. It is the - Thrustxtme (i.e., the impulse) per unit weight of propellant ... 

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