Chapter 2
Conventional chemical propulsion concepts.
2.1 Introduction.
Chemical combustion systems are the most common systems for space application. At least two substances, a fuel and an oxidizer, must be mixed in a specific manner depending on the propellant type and category. Three basic categories exist: liquid propulsion, solid propulsion and hybrid propulsion. The chemical energy associated with the reaction between the fuel and oxidizer is transferred to gaseous products, which expanded through a nozzle produce the thrust for the attached vehicle.
The requirements for thrust level and operating pressure determine the size and the weight of the combustion device:
- Boosters. They are the largest and highest pressure engines. Their high thrust is
required to lift the gross liftoff weight of the entire vehicle from the ground. High pressure is needed mainly to reduce the weight.
- Sustainers or Second-Stage devices. They operate outside of the Earth’s atmosphere
and they can achieve high performance by expanding to very high nozzle exit to throat area ratio.
- Upper-Stage Devices. They are a smaller version of sustainers. Their propellant mass
- Reaction Control and Satellite Propulsion System. They are the smallest rocket thrusters available. These thrusters provide flight vehicle guidance or provide orbit satellite station keeping function.
The following figure illustrates the range of combustion devices of interest in the propulsion community.
Figure 2.1: Liquid rocket engine design application.
2.2 Liquid propulsion system.
A liquid propellant rocket is a rocket with an engine that uses propellants in liquid form [2]. The liquid fuel and the oxidizer are stored in separate tanks, except in the case of monopropellant. The system generally consists of one or more thrust chambers, one or more tanks to store the propellants, a feed mechanism to force the liquids into the thrust chamber, a power source to furnish the energy required by the feed mechanism, suitable plumbing or piping to transfer the liquids, a structure to transmit the thrust forces, and control devices to initiate and regulate the propellant flow rates. Liquid propellants for chemical rockets can be classified as either monopropellants (single substances that are capable, upon igniting, of exothermic break-down) or bipropellants (fuel and oxidizer, substances that are separately stored, pressurized, and then injected into the combustion chamber). In fig. 2.2 are shown some European launcher engines.
Figure 2.2: European launcher engines.
2.2.1 Structure
A liquid rocket engine is composed, as shown in the fig. 2.3, by [3]:
- Tank pressurization (1)
Gas generator or stored gas (N2, He) provide gas to maintain pressure in propellant tank.
A pressurant tank should contain enough gas to completely fill the propellant tank at its desired operating pressure. For stored gas systems, a pressure regulator is set at a desired pressure level; when propellant tank pressure drops below this level, pressurant gas is released into the propellant tank to maintain the desired pressure.
- Propellant storage (2).
Tank pressure must be adequate to allow the desired mass flow rate into the thruster. In case of low gravity or acceleration bladders and other surface tension devices are used.
- Interconnecting plumbing and components (3).
These include feed lines, fill valves, pyrotechnic isolation valves, filters, and heaters. They deliver the propellant to the thruster as efficiently as possible.
Pressure differences or gradients force the flow of propellants from the tanks to the combustion chamber. At the injector, the static pressure of the propellant flow is equal to the chamber pressure. The difference between the static pressure and the total pressure defines the propellant flow rate.
o For monopropellant rockets: tank pressure, without added hardware, feeds
these small systems.
o For bipropellant rockets: in this case two options are possible:1)Tank pressure
feed, in which stored pressure provides propellant flow without additional feed hardware; 2) Pump pressure feed, in which pumps increase the propellant flow pressure, decreasing the required storage pressure in the propellant tank and the tank’s structural mass. For smaller systems, pumps are not usually used. Pumps are typically powered by turbines driven by hot propellant gases. The hot gas source is called the engine cycle (see chapter 2.2.2).
- Flow control. Includes all the smart hardware to control the propellant flow, including
computers and solenoids. We control flow rate by varying flow pressure through the system. For attitude control, the minimum possible thrust impulse is important.
- Thrust chamber (5).
o For monopropellants rockets: propellant feeds into a catalyst bed. The
propellant decomposes into hot gases and is exhausted through a bell nozzle. Most of these systems are radiation cooled. The thruster and flow control valve are usually considered a single, integral component.
o For bipropellant rockets the thrust chamber is composed of:
manifolds: oxidizer dome and fuel manifold supply propellant to the injectors.
injector: injects propellant into combustion chamber. They provide a pressure drop to prevent pressure pulses from moving up into the feed system. Enhances mixing and atomizing of the propellant.
combustion chamber: must have enough volume to allow complete combustion and withstand chamber pressure. Typically cooled.
exhaust nozzle. Allows efficient expansion of exhaust gases. Typically it is a bell nozzle with cooling.
Monopropellant rocket Bipropellant rocket Pressurant tank Pressurant tank Oxidizer tank Propellant tank Valve Nozzle Fuel tank Pump Pump Regulator
Valve Oxidizer dome
Gimbal Combustion chamber Throat (6) Thrust chamber Catalyst bed (1) Tank pressurization (2) Propellant storage (3) Interconnect plumbing and components
(4) Propellant feed system
(5) Flow control
Thrust vector control
Figure 2.3: Scheme of monopropellant end bipropellant rockets.
2.2.2 Engine Cycle or Feed System
The injection system of the propellant is decided by the engine cycles. Four different types are known [1]:
- Pressure fed cycle (see fig. 2.4a).
This type has the simplest feed system. The typical pressurant is gaseous helium or nitrogen. To atomize tank structural weight, pressure-fed flight tank pressures are kept low and both combustion device operating pressures and feed system pressure drops are also minimized. To further minimize pressurant storage bottle and gas weight, the propellant tank pressure may only be regulated over the initial portion of its mission and permitted to operate in a blowdown mode to its propellant exhaustion.
- Gas generator cycle (see fig. 2.4b).
It is the simplest form of the pump fed engine cycles. A fuel-rich gas generator drives turbines of fuel and gas generator cycle of oxidizer turbopump. After expansion in the turbopumps the exhaust gases are dumped into a separate exhaust duct with a small nozzle extension (e.g. Vulcain engine) or the exhaust gases are injected into the nozzle extension at a location where the exhaust pressure matches the pressure in the nozzle extension (Vulcain 2). Because the turbine exhaust gases are dumped overboard at low temperature and low pressure, the lower gas generator exhaust gas performance reduces the overall engine system performance. These systems usually optimize performance at moderate pressure.
- Staged combustion cycle (see fig. 2.4c).
This type of device flows all of one propellant and a small fraction of the other to keep combustion gas temperature low enough to permit turbine drive. Then it injects the remaining propellant downstream of the turbine to recover maximum engine performance at high gas temperature. However, in actual practice, the staged combustion operating pressure is limited from an engine reliability standpoint to a thermal limit to which the combustion device can be cooled. The main combustor in a staged combustion cycle is a gas/liquid injection system, since one propellant circuit has been prevaporized before entering in the turbine.
- Expander cycle (see fig. 2.4d).
It is similar to a staged combustion in which no turbine drive gases are exhausted overboard and furthermore it does not require a preburner or a gas generator. The turbine drive gases are heated while regeneratively cooling the main combustion chamber and nozzle. For the expander bleed cycle (open cycle) the turbine drive gas is dumped into a separate gas duct (as for the gas generator cycle). Except for the expander cycle (closed cycle) all propellants are burnt in the combustion chamber. In practice this cycle is developed only for liquid oxygen/hydrogen propellant combinations because only the liquid hydrogen can provide on adequate cooling. However, its low density requires a system to raise the pump discharge pressure. Expander cycle engines therefore operate at much lower pressures than either staged combustion or gas generator cycle engines and this represents a limitation for use in booster engines.
Figure 2.4: Classification of engine cycles.
Gas Generator Cycle Staged Combustion
Cycle
Expander Cycle
Advantages • Simple interface
between thrust chamber and turbopump • Indipendant development of both subsystems is possible
• Rather low pump pressure required • High chamber pressure s feasible without Isp loss • Compact engine design • No separate turbine exhaust • Most simple engine system • Only CC ignition required • Low temperature turbine • Topping cycle performance
Disadvantages • Isp loss due to
propellant bleed
• Low efficiency
turbines due to low turbine mass flow • Additional
performance loss due thrust chamber mixture shift • Only “integral” development possible due to complex turbopump/thrust chamber interface
• Very high pump
pressures required • Sophisticated engine start procedure • Rather limited chamber pressure (<100bar) • Only “integral” development possible due to complex turbopump/thrust chamber interface • Rather high pump
pressure required Table 2.1: Comparison of different engine cycles.
2.2.3 Propellants
The term liquid propellant includes oxidizer, fuel or the mixture of oxidizer and fuel ingredients.
A monopropellant contains an oxidizing agent and combustible matter in a single substance. Monopropellants must be thermally and chemically stable and it must be easily decomposed and reactive to give good combustion properties.
Hydrazine (N2H4) is an example of monopropellants. It is a toxic, colorless liquid with a high
freezing point. It has had considerable applications in small rockets for altitude control of spacecraft. It is storable for long periods and can be used to provide short-duration thrust.
With exposure to an iridium catalyst it decomposes to ammonia (NH3) and other compounds,
which dissociate at high temperature. At elevated temperatures many materials decompose hydrazine, including iron, nickel and cobalt. Hydrazine is characterized by a positive heat of formation and, therefore, generally gives good performance when compared with many common fuels. At a combustion chamber pressure of 6.94 MPa it can theoretically generate a specific impulse of 245s.
A bipropellant rocket unit has two separate propellants, an oxidizer and a fuel. Table 2.2 and Table 2.3 give the most common used. Propellants are stored separately and are not mixed outside the rocket engine. The majority of liquid propellant rockets have been manufactured for bipropellant application.
Cryogenic propellants are liquefied gases at low temperature, such as liquid oxygen (-147°C) or liquid hydrogen (-253°C). The oxygen-hydrogen combination has now been the subject of great experience.
Many different types of storable and cryogenic liquid oxidizer propellants have been proposed like boron-oxygen-fluorine compounds, oxygen-fluorine compounds, nitrogen-fluorine formulations, and fluorinated hydrocarbons. They have high specific impulse, but also undesirable characteristics.
Only few liquid oxidizers are used commonly today:
- Liquid oxygen (O2). The boiling point is at 90K at atmospheric pressure; it has a
specific gravity of 1.14 kg/cm3 and a heat of vaporization of 213 kJ/kg. It has been
used in combination with alcohols, jet fuels (like kerosene type), gasoline and hydrogen. Liquid oxygen was used in various missiles and vehicles: with jet fuel in Atlas, Thor, Jupiter, Titan I, Saturn booster; with hydrogen in Space Shuttle and Centaur upper stage; with 75% alcohol in Redstone. Liquid oxygen is a non toxic and non corrosive liquid and will not appreciably deteriorate the container walls.
- Nitric acid (HNO3). There are several types of nitric acid mixtures, the most common
is red fuming nitric acid, which consists of concentrated nitric acid that contains between 5 and 20% dissolved nitrogen dioxide. It is highly corrosive. A small addition of fluorine ion inhibits the nitric acid and reduces the corrosion with many metals. Lime and alkali metal hydroxides and carbonates are common neutralizing agents. It has been used with gasoline, various amines, hydrazine and alcohols. Its specific gravity varies from 1.5 to 1.6, depending on the percentage of water, nitric oxide and impurities.
- Hydrogen Peroxide (H2O2). In rocket application, hydrogen peroxide has been used in
a highly concentration form of 70 with water. Because of its storage stability
problems, it is not very used. In the combustion chamber, H2O2 decomposes forming
superheated steam and gaseous oxygen. This decomposition is brought about by the action of catalysts such as various permanganates, manganese dioxide, platinum, and iron oxide. The theoretical specific impulse of 90% hydrogen peroxide is 147s, when used as a monopropellant.
- Nitrogen Tetroxide (N2O4). It is the most common storable oxidizer used today, its
liquid temperature range is narrow and it is easily frozen or vaporized. It can be stored indefinitely in sealed containers made of compatible material. It is hypergolic with many fuels and can cause spontaneous ignition with many common materials. Nitrogen tetroxide is used in Titan II engines together with a fuel mixture consisting of
hydrazine and unsymmetrical dimethylhydrazine. It is also used with
monomethylhydrazine fuel in the Space Shuttle orbital maneuver system and reaction control system and in many spacecraft propulsion systems like Aestus.
Many different chemicals have been proposed, investigated, and tested, but only a few have been used in production rocket engines. the most common are:
- Hydrocarbon Fuels. Most of them are in use with other application and engines, as
gasoline, kerosene, diesel oil, and turbojet fuel. Their physical properties and chemical composition vary widely with the type of crude oil from which they were refined, with the chemical process used in their production, and with the accuracy of control exercised in their manufacture. They are relatively easy to handle, and there is an
ample supply of these fuels available at low cost. The most hydrocarbon fuel used is the RP-1 in the combination with liquid oxygen.
- Liquid Hydrogen (H2). It is the lightest fuel, having a specific gravity of 0.07 and a
boiling point of about 20K. The low fuel density requires very bulky fuel tanks, which necessitate very large vehicle volumes. The extremely low temperature makes the problem of choosing suitable tank and piping materials difficult, because many metals lose their strength at low temperatures. Liquid hydrogen is used with liquid oxygen as an oxidizer in Ariane 5 (fig.2.5).
Table 2.2: Classification and characteristics of liquid oxidizers.
2.2.4 Advantages and disadvantages
Liquid propellant rockets have many advantages:
- they can be throttled and have a good control of mixture ratio; they can also be shut
down, and, with a suitable ignition system or self-igniting propellant, restarted.
- they have high performance that usually corresponds to high specific impulse.
- a liquid rocket engine (LRE) can be tested prior to use, whereas for a solid rocket
motor a rigorous quality management must be applied during manufacturing to insure high reliability.
- a LRE may be reused for several flights, like in the Space Shuttle.
Use of liquid propellants can be associated with a number of issues:
- Because the propellant is a very large proportion of the mass of the vehicle, the center
of mass shifts significantly rearward as the propellant is used; the control of the rocket is typically lost if its center mass gets too close to the center of drag.
- Liquid propellants are subject to slosh, which has frequently led to loss of control of
the vehicle. This can be controlled with slosh baffles in the tanks as well as judicious control laws in the guidance system.
- Liquid propellants can leak, especially hydrogen, possibly leading to the formation of
an explosive mixture.
- Non-storable liquid rockets require considerable preparation immediately before
launch. This makes them less practical than solid rockets for most military application.
A comparison between the different types of liquid propellants is given in the following table:
Advantages Disadvantages
Storable propellants
• storability for long duration missions
• higher density
• limited energy content • more expensive • more corrosive
Cryogenic • maximum specific impulse
• economical
• evaporative losses • schedule and logistics
Hypergolic propellants
• unlimited restart capability • capable of pulse mode
operation
• toxic
• more corrosive • more expensive
Non hypergolic propellants • less toxic
• more economical
• need external source of ignition
• chamber pressure
Monopropellants
• simplicity of system
reliability
• capable of very short pulse widths
• relatively low specific impulse
• limited catalyst life
Storable bipropellants
• higher specific impulse
• higher total impulse
capability
• complexity of dual feed system
• must maintain mixture ratio to avoid residuals
• corrosive oxidizers
• very short pulses are off limits for space shuttle • system inoperative once
one component leaks out Table2.4: Comparison of different types of propellants.
2.3 Solid propulsion system
In solid propellant rocket motors the burned propellant is placed in the combustion chamber or case. Solid propellant charge is called a grain and it contains all the chemical elements for complete burning. Once ignited, it usually burns smoothly on the exposed surface of the charge. Because there are no feed system or valves, solid propellant rocket engine are usually relatively simple in construction (see fig. 2.6).
Figure 2.5: Ariane 5 is a large two stage liquid propellant system, 59 m of height and 5.4 m of
2.3.1 Components and subsystems
A solid rocket engine is composed by (see fig. 2.7):
- Thrust skirt (1).
It connects the motor case to vehicle structure and transfers thrust load to the vehicle.
- Motor case (2).
Usually it is made of titanium, high-strength steels or wound fiber. It contains combustion pressure.
- Polar boss (3).
It is used to transfer nozzle loads to motor case and it connects nozzle and case.
- Nozzle assembly (4).
It controls expansion of chamber gases. It is made of high temperature material to minimize heat and erosion problems. Typical materials are carbon/phenolic, silica-phenolic, or carbon/carbon. The thrust vector controls vary the nozzle angle. the characteristic parameters are expansion ratio, cone angle, throat area and the exit area.
- Igniter (5).
It provides enough energy to start the solid propellant combustion.
- Internal insulation (6).
It is used to protect the motor case from the high propellant temperatures. In fact it is made of ablative material to dissipate heat and its low thermal conductivity reduces heat transfer to motor structure.
- Propellant grain (7).
It contains granular fuel and oxidizer in a rubber matrix binder. Typical fuels are metal powders such as aluminium and typical oxidizers are ammonium perchlorate or ammonium nitrate. A typical binder is polybutadiene
- Port or bore (8).
It provides the propellant surface area for combustion; more surface area drives up chamber pressure and thrust level. A motor can have many numbers of ports, depending on thrust and pressure requirements because the port shape determines burn surface area over the burn duration. Typical shapes are star patterns and cylinders with transverse slots.
Figure 2.7: Scheme of solid motor rocket
2.3.2 Solid propellants
Three types of solid propellant are distinguished, according to the distribution of the fuel and the oxidant.
If the fuel and oxidant are contained within the same molecule, the propellant is called homogeneous. The ingredients are mixed at molecular level. Double base propellants, combination of nitroglycerin-nitrocellulose with small quantities of additives are the most common used. They are usually used only for low cost military or civilian applications.
Heterogeneous mixtures oxidizing crystals in an organic plastic like fuel binder are called composite propellants. The ingredients are mixed only mechanically and thus distinguishable on macroscopic scale. The main component is an inorganic crystalline salt, an oxidizer dispersed in a polymeric binder acting also as fuel. Often, but not always, metal powders added as high energy fuel for high performance applications.
The oxidizers usually consist of ground crystals and most used are: ammonium perchlorate
(NH4ClO4 in a short form AP), ammonium nitrate (NH4NO3 in a short form AN), potassium
perchlorate (KClO4 in a short form KP) for pyrotechnic igniters, potassium nitrate (KNO3 in a
withstand severe thermal and mechanical stress without cracking. Binders are mainly made of synthetic rubbers. the most used are polybutadiene based rubbers such as
Hydroxyl-Terminated Polybutadiene (HTPB), Carboxy-Terminated Polybutadiene (CTPB),
Polybutadiene Acrylic Acid (PBAA).
Based on oxidizer grain distribution, composite solid propellants are distinguished as monomodal (used only for a lab purposes), bimodal (common), trimodal (common), and tetramodal (rare).
Composite modified Double base (CMDB) or mixed formulations are mixture methods and they are mainly used for high performance military applications.
2.3.3 Propellant burning rate and internal ballistic
Energetic materials in general are capable of a dual reacting regime: the supersonic regime and the subsonic regime. In the supersonic regime a combustion wave preceded by a strong shock wave brings about a detonation wave, propagating at a speed of several km/s and limited by the thermo chemical energy content of the reacting material; in the subsonic regime a combustion wave brings about a deflagration wave, propagating at a speed of cm/s and limited by heat and/or mass diffusion. Deflagration waves in general consist of an initial condensed phase and a final phase. Burning surface is the interface between condensed phase and phase. The propagation rate of this interface is called burning rate; it can also be seen as the regression rate of the condensed phase. There are applications that are convenient to define a linear burning rate as the web thickness burned per unit time in the direction perpendicular to the burning surface.
The design and operation of solid rocket motors depend on the combustion features of the propellant charge (burning rate, burning surface, and grain geometry) and their evolution in time.
The burning surface of a propellant grain propagates in a direction perpendicular to the surface. Success in rocket motor design and development depends significantly on knowledge of burning rate behavior 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 by:
- details of chemical composition (catalyst, modifiers, additives, usually present in a small or fractional percentage)
- physical effects (particle size distribution, presence of wires or staples)
- details of manufacturing process and other miscellaneous factors
- operating conditions (pressure, initial temperature, natural or external radiation, heat
losses, gas flow parallel to the burning surface, acceleration)
- mode of operation (steady or unsteady)
2.3.4 Propellant grain and grain configuration
Grain is the shaped mass of propellant inside the rocket motor. The propellant material and the geometrical configuration of grain determine the motor performance characteristic. Propellant grain is a cast, molded or extruded body and its appearance and feel is similar to hard rubber or plastic. Grain must maintain structural integrity during shipment, storage, and rocket operation.
As a result of motor developments of the past three decades, many grain configurations are available to motor designers. The different configurations are used as thrust control during the propulsive mission (see fig. 2.8).
2.3.5 Advantages and disadvantages
A typical solid propellant rocket may have a specific impulse between 200 to 300 s thus it has relatively low specific impulse compared to the liquid rocket engine.
For their low cost, solid engines have been used as initial stages in rockets (the classic example is the Space Shuttle or Ariane family rockets), while reserving high specific impulse engines, especially less massive hydrogen fuelled engines for higher stages. The extreme difficulty of varying thrust on demand limits the application of rocket engines. In addition, they are used as the final boost stage for satellites due to their simplicity, reliability, compactness and reasonably high mass fraction. An attractive attribute for military use is the ability to remain loaded in the rocket for long durations and then reliably launch at the moment’s notice, it is due to their chemical and physical stability (long service life, around 15 years).
2.4 Comparison with hybrid rockets
Although the liquid rocket is an efficient high performance system, it is quite complex and requires costly plumbing. The solid engine avoids this problem by premixing the fuel and the oxidizer in a solid form that is dense and compact. The major disadvantages, however, are the explosive danger and the lack of thrust control or termination. Although the propulsion system per se may be cheaper, these latter disadvantages translate to a more costly vehicle design to accommodate such requirements as acceleration, atmospheric heating and shutdown. The hybrid rocket [5], with half the plumbing of the liquid, but retaining its operational flexibility and avoiding the explosiveness of the solid, therefore, provides an attractive alternative option.
