Yıldız kesitli katı yakıcı-yakıtlı bir roket motorunun performansının incelenmesi
A study of performance of a star-shaped solid propellant rocket motor
Yazar: HALİT YAŞAR
Danışman: DOÇ.DR. VELİ ÇELİK
İstanbul Teknik Üniversitesi / Fen Bilimleri Enstitüsü
Derece: Yüksek Lisans, 1992
Konu: Makine Mühendisliği = Mechanical Engineering
Dizin: Katı yakıtlar = Solid fuels ; Roket motorları = Rocket engines

Özet:
ÖZET Bu çalışmada, yıldız kesitli çekirdeğe C grain} sahip katı yakıcı -yakıtlı bir roket motorunun performansı, ideal roket varsayımına göre düzenlenen bir bilgisayar programı yardımıyla hesaplanmakta ve elde edilen ideal performans değerleri, aynı model roket motorunun deneysel performans değerleriyle karşılaştırılmaktadır. Çalışmanın ilk bölümünde katı yakıcı -yakıtlı roket motorlarıyla ilgili kısa ve genel bilgiler verilmektedir. İkinci bölümde, bir katı yakıcı-yakıtlı roket motorunun temel parçaları tanıtılmakta ve katı yakıcı -yakıtlar hakkında bilgi verilmektedir. Üçüncü bölümde, katı yakıcı -yakıtların yanma mekanizmaları, yanma hızı ve yanma hızı ölçüm teknikleri anlatılmaktadır. Dördüncü bölümde, ideal roket, izantropik lüle akı sı ve performansla ilgili ifadeler verilmektedir. Beşinci bölümde, geliştirilen ideal performans hesap lama programı tanıtılmakta, yıldız kesitli çekirdeklerde yanma yüzeyinin hesaplanmasına ilişkin matematiksel model verilmekte ve elde edilen deneysel değerlerle nümerik çözüm sonuçlarının karşılaştırılması yapılmaktadır. Son bölümde ise, elde edilen sonuçlar özetlenmekte ve bu sonuçlarla ilgili düşünceler belirtilmektedir. Çalışmanın sonucunda, elde edilen ideal performans değerleriyle deneysel sonuçlar arasında oldukça iyi bir uyum olduğu gözlemlenmiştir. Değerler arasındaki farklılaşmalar literatürde belirtilen sınırlar içinde kalmaktadır.

Abstract:
A STUDY OF PERFORMANCE OF A STAR-SHAPED SOLID PROPELLANT ROCKET MOTOR SUMMARY In general, a vehicle is propelled by forces, called thrusts, which provide a desired component of acceleration. These forces can be produced in a variety of ways. Solid propellant rockets are examples of a pure reaction system in which the propulsive forces are produced by the ejection of mass(propellants) initially contained in the system. Self-contained systems of this type are called rocket motors and can operate in space as well as in atmospheres, since they do not require an external propulsive fluid. Four categories of rocket motors may be defined, according to the physical state of the propellant materials carried within the rocket. These are; solid propellant rocket motors, liquid propellant rocket motors, gaseous propellant rocket motors and hybrid rocket motors (which contain propellants stored in at least two of the three of physical states of matter). This study is only concerned with solid propellant rocket motors. A solid propellant rocket is the simplest form of chemical propulsion. The fuel and oxidizer are both incorporated in a single solid, called the propellant grain, located inside a container called the combustion chamber. This is large in comporation with the combustion chamber of a liquid propellant rocket motor. A device called as igniter, which is designed to initiate the burning, is placed inside the central cavity of the combustion chamber. After ignition, the hot gases, which are produced when the solid burns, flow through the central cavity and are accelerated to a high velocity by means of a nozzle. During combustion, gases evolve from the solid propellant grain only at its surface. Thus the surface of the solid regresses normal to itself during burning, at the "linear regression rate" of the propellant grain. The combustion gases come into contactwith the outer Vllshell or case o the chamber only at the end of the burning. A solid propellant rocket motor is composed of four basic parts: -the propellant grain, -the case, -the nozzle, -the igniter. The grain geometry provides a basis for a preliminary classification of solid rocket propulsion systems. Once the grain is ignited, burning generaly progresses until the propellant is completely consumed. The burning surface geometry and its time evolution will then impose the thrust-time history. The grain is called neutral if the thrust remains constant during the entire burning history. This occurs when the total burning surface area does not vary with time. There also exist progressive grains for which the thrust increase with time and regressive grains for which the thrust decrease with time. Internally burning cylinders or tubes are progressive, while rod-shaped grains are regressive. The most common current configurations are star-shaped cylindrical grains which provide a relatively large propellant surface area for burnt gas emission while maintaining approximately neutral burning. The geometry of the motor case is related to that of the grain. Case design also depends on the application. Two principal types of casing materials are currently employed, metalic materials and glassy materials. The nozzle attached to the downstream end of the rocket motor consists of a convergent section, a narrow-diameter throat and a divergent section. Since solid propellant rocket nozzles are generally uncooled, it is necessary to use nozzle materials capable of withstanding a high thermal load. Solid propellant rocket igniters often consist of electrically initiated, conventional pyrotechnic compositions. The combustion of an auxiliary propellant contained in the igniter generates hot gases which come into contact with the grain surface and induce ignition of the grain. Solid propellants are chemicals, which produce hot, high pressure gases by means of a combustion process, and these gases can therefore be used to provide the reaction force for rocket propulsion. Solid propellants usually vixiinclude two or more of the following species: 1. Oxidizer {nitrates or perclorates), 2. Fuel (organic resins or plastics), 3. Chemical compount combining fuel and oxidizer qualities (nitrocellulose or nitroglycerin) 4. Additives (to control fabrication process, burning rate, etc.) 5. Inhibitors (bended, tabed, or dip-dried onto propellant) to restrict burning surface. There are two principal types of propellants. The first, often called composite propellant type, has two important ingredients, a fuel and an oxidizer, neither of which would burn satisfactorily without presence of the other. Often these concist of crystaline, finely ground oxidizers dispersed in a matrix of a fuel compound. The second type contains unstable chemical compounds, such as nitrocellulose or nitroglycerin, which are capable of combustion in the absence of all other materialls. Because many of these propellants are based largely on a colloid of nitrocellulose and nitroglycerin, they are often referred to as double-base propellants. Most of the accepted solid propellants contain from four to eight different chemicals. In addition to the principal ingredients, small percentages of additives are used to control the physical and chemical properties of the solid propellant. The analysis of ideal rockets assumes one-dimensional flow. The usefulness of this concept is indicated by the fact that the measured performance is usually within 1 to 10 % of the ideal values. In designing new rockets, it has become accepted practice to use ideal rocket parameters which are then modified by appropriate corrections. An ideal rocket unit is one in which the following assumptions are valid: l.The working substance (propellant products) is homogeneous and invariant in composition throughout the rocket chamber and nozzle. 2. The working substance obeys the perfect gas laws. 3. There is no friction. 4. There is no heat transfer across the rocket walls; therefore the flow is adiabatic. IX5. The propellant flow is steady and constant. The expansion of the working fluid takes places in a uniform and steady manner without shock, vibration, or discontinuities. Transient effects (start, stop) are neglected. 6. All the exhaust gases leaving the rocket nozzle have an axially directed velocity. 7. The gas velocity, pressure, temperature, or density is uniform across any section normal to the nozzle axis. 8. Chemical equilibrium is established within the rocket chamber and does not shift in the nozzle. 9. All the species of the working fluid are gaseous. Any condensed phases (liquid or solid) have negligible mass. Because the chamber temperatures are high, the gases are well above their respective saturation and follow the perfect gas laws very closely. Postulating no friction and a steady flow without heat transfer to the wall allows the use of isentropic expansion relations in the rocket nozzle, thereby permitting the assumption of a maximum conversion of heat energy into the kinetic energy of the jet. This implies that the nozzle flow is thermodynamic ally reversible. The burning surface of a propellant grain recedes in a direction essentially perpendicular to the surface. The rate of regression, expressed in m/sn, is the burning rate r. Success in rocket motor design and development depends significantly on knowledge of burning rate behaviour of the selected propellant under average motor operating conditions and design limit conditions. Burning rate is a function of the propellant composition itself and can be increased in the following ways: l.Add a burning rate catalyst, or increase percentage of catalyst. 2. Decrease the oxidizer particle size. 3. Increase oxidizer percentage. 4. Increase the binder heat of combustion. 5. Imbed wires or metal staples in the propellant. Composite propellants can be influenced by all five. Item 2 and 3 are not applicable to double-base propellants. Emprical burning-rate data are usually obtained from tests using the standart (Crawford) Strand-burner, where a rod or "strand" of propellant is burned from one end to the other, or small-scaleballistic evaluation motors; they are preferred for motor design purposes. Strand-burner data is useful in screening propellant formulations and in quality control operations. Data from full-scale motors tested under a variety of conditions constitute the final proof of burning-rate behavior. Aside from the propellant formulation and manufacturing process, burning-rate in a full-scale motor is influenced by the following: 1. Combustion chamber pressure, 2. Initial temperature of the propellant, 3. Combustion gas temperature, 4. Velocity of the gas flow parallel to the burning surface, 5. Motor motion (acceleration and spin-induced grain stress). The burning rate of propellant in a rocket motor is a function of many parameters, and at any instant governs the mass flow rate of hot gas generated and flowing from the motor. With many propellant it is possible to approximate the burning rates as a function of chamber pressures, at least over a limited range of chamber pressures. In many production-type propellants, whether double-base, composite, or composite double-base, this emprical equation is, r = a + b.Pn where r, the burning rate, is usually cm per second, and the chamber pressure is atm, a and b are emprical constant influenced by ambient grain temperature, and n is known as the burning rate pressure exponent. The burning rate pressure exponent must be less than unity for stable combustion, as may be demonstrated by the above simple argument. Neglecting the relatively small gas storage terms, nozzle flow rate and gas generation rate must be equal. Nozzle gas flow rate is directly proportional to P. Concider a small decrease from the operating point pressure. It can be seen that for n l any small decrease in chamber pressure means that nozzle flow rate will exceed gas generation rate and the pressure will drop even further. Thus for n>l, small disturbances are amplified and the combustion can not be xistable. Hence for stable combustion, n must be less than unity. Those parameters that govern the burning rate and mass discharge rate of motors are called internal ballistic properties; they include r, K,