Considering the complexities of the liquid propellant rocket engine, it does not seem remarkable that so much attention has been given to the design and development of the much simpler solid propellant motor. This has a range of applications: the main propulsion system for small and medium launchers; as a simple and reliable third stage for orbital injection; and most of all as a strap-on booster for many modern heavy launchers. The solid propellant is storable, and is relatively safe to handle; no propellant delivery system is required, and this produces a huge improvement in reliability and cost. There are two main disadvantages: the motor cannot be controlled once ignited (although the thrust profile can be preset), and the specific impulse is rather low because of the low chemical energy of the solid propellant.
Thermodynamically a solid-fuelled rocket motor is identical to a liquid-fuelled engine. The hot gas produced by combustion is converted to a high-speed exhaust stream in exactly the same way, and so the nozzle, the throat and the restriction in the combustion chamber leading to the throat are all identical in form and function. The thrust coefficient is calculated in the same way as for a liquid-fuelled engine, as is the characteristic velocity. The theoretical treatment in Chapter 2 serves for both.
The hot gas is produced by combustion on the hollow surface of the solid fuel block, known as the charge, or grain. In most cases the grain is bonded to the wall of the combustion chamber to prevent access of the hot combustion gases to any surface of the grain not intended to burn, and to prevent heat damage to the combustion chamber walls. The grain contains both fuel and oxidant in a finely divided powder form, mixed together and held by a binder material.
Figure 4.1 shows a typical solid-motor configuration. In comparison with the liquid rocket combustion chamber it is very simple. It consists of a casing for the propellant, which joins to a nozzle of identical geometry to that of a liquid-fuelled
engine. Once the inner surface of the grain is ignited, the motor produces thrust continuously until the propellant is exhausted.
The fundamental simplicity of the solid propellant rocket enables wide application. The exhaust velocity is not very high—the most advanced types can produce about 2,700 ms"1—but the absence of turbo-pumps and separate propellant tanks, and the complete absence of complicated valves and pipelines, can produce a high mass ratio, low cost, or both. In addition, the reliability is very high, due to the small number of individual components compared with a liquid-fuelled engine. The one big disadvantage is that the device cannot be test fired, and so the reliability has to be established by analogy and by quality control. The two areas in which solid motors excel are as strap-on boosters and as upper stages, particularly for orbit insertion or for circularisation of elliptical transfer orbits. Solid propellants are, by definition, storable.
As a booster, a solid motor can have a very high mass-flow rate and therefore high thrust, while the engineering complexity and cost can be low in a single use item. This is ideal for the early stages of a launch where high exhaust velocity is not an issue. To produce the same thrust with a liquid-fuelled rocket would not require such a large engine, because of the higher specific chemical energy of some liquid propellants, but it would be much more costly and less reliable. Very large solid boosters can be made and fuelled in sections which are then bolted together, which again makes for simplicity of construction and storage of what would otherwise be a very large unit.
As a final stage the solid motor is again reliable, and is well adapted to high massratio. While the dead weight of a liquid stage includes turbo-pumps and empty tanks for two separate propellants, the dead weight of a solid stage is just the casing and the nozzle. The casing for upper stages is often made of composite materials, reducing the mass even further. It is also convenient to make such a stage with a spherical or quasispherical form, so as to minimise the mass of containing walls.
In comparison with a liquid-fuelled engine, the solid motor is very simple, and the design issues are therefore fewer. There is no injector, and no propellant distribution system. Design issues related to the propellant are mostly concerned with selection of the propellant type and the mounting and protection of the propellant in the casing, and ignition is similar to that of a liquid-fuelled engine. There are no propellant tanks, but the casing has to contain the propellant and also behave as a combustion chamber. For boosters the casing is large, and to combine large size with resistance to high combustion pressure is very different from the same issue in a liquid system where the requirements are separated. Cooling is totally different, because there are no liquids involved and heat dissipation has to be entirely passive.
Combustion stability—which for a liquid-fuelled rocket is dependent only on a steady supply of propellant once the chamber and injector have been optimised—is very complicated for a solid propellant. Here the supply of combustible material is dependent on conditions in the combustion chamber, and there are increased chances for instabilities to arise and propagate. Associated with stability is thrust control. For a liquid rocket the thrust is actively controlled by the rate of supply of propellants, and in the majority of cases it is stabilised at a constant value. For a solid rocket the thrust depends on the rate of supply of combustible propellant; this depends on the pressure and temperature at the burning surface, and it cannot actively be controlled. In the same way, a liquid-fuelled engine can be shut down by closing valves, whereas the solid motor continues to thrust until all the propellant is exhausted. These design problems are central to the correct performance of a solid propellant rocket motor.
While the solid-fuelled rocket is essentially a single-use item, the cost of large boosters is very high, and the necessary engineering quality of some components— specifically the casing—may make them suitable for reuse. This was a design feature for both the Space Shuttle and the Ariane 5 solid boosters. The Space Shuttle boosters are recovered, and the segments are reused. The Ariane 5 boosters are also recovered, but only for post-flight inspection.
While there is a wide choice of propellant composition for liquid-fuelled rocket engines, the choice is considerably more narrow for solid propellants. Rather than selecting a particular propellant for a particular purpose, each manufacturer has its own optimised propellant mixture. The basic sold propellant consists of two or more chemical components which react together to produce heat and gaseous products. Solid propellants have been used since the earliest times, and until this twentieth century were based on gunpowder—a mixture of charcoal, sulphur and saltpetre. Modern propellants do not differ in fundamentals from these early mixtures. The oxidant is usually one of the inorganic salts such as potassium nitrate (saltpetre) although chlorates and perchlorates are now more commonly used. The fuels sometimes include sulphur, and carbon is present in the form of the organic binder.
As with any other type of rocket, the aim is to achieve the highest combustion temperature together with the lowest molecular weight in the exhaust. The difficulty with solid oxidants is that they are mostly inorganic and contain metal atoms. These lead to higher molecular weight molecules in the exhaust. Similarly the solid fuels generally have a higher atomic weight than hydrogen, and so again the molecular weight in the exhaust is driven up. The chemical energy, per unit mass of propellant, can be the same as for the main liquid propellants, and so the combustion temperature is similar. A particular problem is that some of the combustion products may form solid particles at exhaust temperatures. This affects the performance of the nozzle in converting heat energy into gas flow. All of these properties affect the performance of solid motors.
The charge of propellant in a solid rocket motor is often called the grain. The basic components of the grain are fuel, oxidant, binder, and additives to achieve burning stability and stability in storage. The finished charge must also be strong enough to resist the forces induced by vehicle motion and thrust. It must also be thermally insulating to prevent parts of the grain—other than the burning surface— from reaching ignition temperature.
In the past, two different kinds of solid propellant have been used. The first kind is the mixture of inorganic oxidants with fuels, as described above. This is the most commonly used today. The other type is based on nitrated organic substances such as nitroglycerine and nitrocellulose. These came into use as gun propellants after gunpowder, and it was natural that they should be considered as rocket propellants. These materials have the property that they contain the oxidant and fuel together in a single molecule or group of molecules. Heat induces a reaction in which the complex organic molecule breaks down, which produces heat and gaseous oxides of nitrogen, carbon and hydrogen. The molecular weight of such gas mixtures is rather low, giving an advantage in terms of exhaust velocity. These propellants are termed homogeneous propellants, for obvious reasons. They are not used for launcher boosters and most orbital change motors, because they have been superseded by more advanced mixture propellants.
The fundamental requirement is to develop high thrust per unit mass. As discussed in Chapter 2, this requires a high combustion temperature and low molecular weight of the combustion products. In general a relatively high temperature of combustion is easy to achieve, but it is impossible to have the same low molecular weight of the products achievable with liquid hydrogen and liquid oxygen. The presence of carbon and the byproducts of the inorganic oxidants, potassium and sodium salts, produces a higher molecular weight and hence a lower exhaust velocity. Referring back to Chapters 1 and 2, we can see that high molecular weight does not prevent the solid motor from developing high thrust, which is just a matter of high mass flow and throat area. High ultimate vehicle velocity is harder to achieve with a solid motor because of the low exhaust velocity. A typical value would be about 2,700 ms_1. For final stages, optimisation is directed towards improving the mass ratio rather than the exhaust velocity.
In modern propellants metallic powders are often added to increase the energy release and hence the combustion temperature. Aluminium is usual, and in this case the exhaust products will contain aluminium oxide, which has a high molecular weight and is refractory, and so is in the form of small solid particles. Particles in the exhaust stream reduce efficiency: they travel more slowly than the surrounding highvelocity gas, and they radiate heat more effectively (as black bodies) and therefore reduce the energy in the stream. The loss of exhaust velocity may be balanced by the higher combustion temperature and an increase in effective density of the exhaust gases. This increases the mass flow and hence the thrust. High thrust is applicable for a first-stage booster where ultimate velocity is not as important as the thrust at liftoff. In designing a motor for high thrust, increasing the exhaust density may be preferable to an increase in throat diameter and hence in overall size of the booster; the mass ratio is also increased if the grain density is higher. The presence of particles in the exhaust produces the characteristic dense white 'smoke' seen when the boosters ignite. The exhaust from a liquid-fuelled engine is usually transparent.
The most commonly used modern solid propellant is based on a polybutadiene synthetic rubber binder, with ammonium perchlorate as the oxidiser, and some 1216% of aluminium powder. The boosters for the Space Shuttle use this type of propellant, as do the boosters for Ariane 5 and many upper stages. The combustion temperature without the aluminium is about 3,000 K with 90% of ammonium perchlorate. The addition of 16-18% aluminium increases the temperature to
3,600 K for the Ariane 5 booster, and the oxidiser concentration is reduced correspondingly.
The chemical composition of the exhaust is approximately 32% aluminium oxide, 20% carbon monoxide, 16% water, 12% hydrogen chloride, 10% nitrogen, 7% carbon dioxide and 3% chlorine and hydrogen. A major part of the aluminium oxide condenses into solid particles, but fortunately this does not contribute to the molecular weight in the expanding gases: Al203 has a molecular weight of 102. The combined effect of the gaseous components is to produce an average molecular weight of about 25. The combustion parameter is 12, giving a characteristic velocity of 1,700ms"1. The particles will reduce the mean exhaust velocity because of the effects mentioned above. The quoted vacuum exhaust velocity is 2,700 ms"', which is fairly close to the theoretical value if we assume a reasonable thrust coefficient. So this two-phase flow—in which the exhaust gases follow the normal expansion, cooling and acceleration, alongside particles which are accelerated by the gas— does not reduce the exhaust velocity very much. If the particles were to evaporate then a very high molecular weight gas would result, producing a very low exhaust velocity. This solid propellant is therefore rather efficient in producing high thrust and a reasonable exhaust velocity.
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