固体火箭设计

Solid-Fueled Rocket
When the fuel in a solid-fueled rocket is ignited, the gases formed during combustion are forced out the nozzle and the rocket moves forward. The fuel is called the grain and is often formed with a hollow core for longer burning times.
Solid rockets are rockets with a motor that uses solid propellants (fuel/oxidizer). The Chinese invented solid rockets and were using them in warfare by the 13th century. All rockets used some form of solid or powdered propellant up until the 20th century. Solid rockets are considered to be safe and reliable due to the long engineering history and simple design.
Basic Concepts
simple solid rocket motor consists of a casing, nozzle, grain (propellant charge), and igniter.
The grain behaves like a solid mass, burning in a predictable fashion and producing exhaust gases. The nozzle dimensions are calculated to maintain a design chamber pressure, while producing thrust from the exhaust gases.
Once ignited, a solid rocket motor cannot be shut off.
Modern designs may also include; steerable nozzle for guidance, avionics, recovery hardware (parachutes), self destruct mechanisms, APU's, and thermal management materials.
Design
Design begins with the total impulse required, this determines the fuel/oxidizer mass. Grain geometry and chemistry are then chosen to satisfy the required motor characteristics.
The following are chosen or solved simultaneously. The results are exact dimensions for grain, nozzle and case geometries;
•The grain burns at a predictable rate, given its surface area and chamber pressure.
•The chamber pressure is determined by the nozzle orifice diameter and grain burn rate.
•Allowable chamber pressure is a function of casing design.
•The length of burn time is determined by the grain 'web thickness'.
The grain may be bonded to the casing, or not. Case bonded motors are much more difficult to design, since deformation of both the case and grain, under operating conditions, must be compatible.
Common modes of failure in solid rocket motors are; fracture of the grain, failure of case bonding, and air pockets in the grain. All of these produce an instantaneous increase in burn surface area, and a corresponding increase in exhaust gas and pressure, and rupture of the casing.
Another failure mode is casing seal design. Seals are required in casings that have to be opened to load the grain. Once a seal fails, hot gas will erode the escape path and result in failure. This was the cause of the Space Shuttle Challenger disaster.
Grain
Solid fuel grains are usually molded from a thermoset elastomer (which doubles as fuel), additional fuel, oxidizer, and catalyst. HTPB is commonly used for this purpose. Ammonium perchlorate is the most common oxidizer used today.
The fuel is cast in different forms for different purposes. Slow, long burning rockets have a cylinder shaped grain, burning from one end to the other. Most grains, however, are cast with a hollow cross section, burning from the inside out (and outside in, if not case bonded), as well as from the ends.
The thrust profile over time can be controlled by grain geometry. For example, a star shaped hole down the center of the grain will have greater initial thrust because of the additional surface area. As the star points are burned up, the surface area and thrust are reduced.
Casing
The casing may be constructed from a range of materials. Cardboard is used for model engines. Steel is used for the space shuttle boosters. Filament wound graphite epoxy casings are used for high performance motors.
Nozzle
A Convergent Divergent design accelerates the exhaust gas out of the nozzle to produce thrust.
Sophisticated solid rocket motors use steerable nozzles for rocket control. Performance
Solid fuel rocket motors have a typical specific impulse of 265 lbf·s/lb (2.6 kN·s/kg). This compares to 285 lbf·s/lb (2.8 kN·s/kg) for kerosene/Lox and ~389 lbf·s/lb (3.8 kN·s/kg) for liquid hydrogen/Lox1. For this reason solids are generally used as initial stages in a rocket, with better performing liquid engines reserved for final stages. However, the venerable Star line motors manufactured by Thiokol have a long history as the final boost stage for satellites. This is due to their simplicity, compactness and high mass fraction. The ability of solid rockets to remain in storage for long periods, and then reliably launch at a moments notice, makes them the design of choice for military applications.
Amateur rocketry
Solid fuel rockets can be bought for use in model rocketry; they are normally small cylinders of fuel with an integral nozzle and a small charge that is set off when the fuel is exhausted. This charge can be used to ignite a second stage, trigger a camera, or deploy a parachute.
Designing solid rocket motors is particularly interesting to amateur rocketry enthusiasts. The design is simple, materials are inexpensive and constructions techniques are safe.
Early amateur motors were gunpowder. Later, zinc/sulfur formulations were popular.
Typical amateur formulations in use today are; sugar (sucrose, dextrose, and sorbitol are all common)/potassium nitrate, HTPB (a rubber like epoxy)/magnesium/ammonium nitrate, and HTPB or PBAN/aluminum/ammonium perchlorate. Most formulations also include burn rate modifiers and other additives, and also possibly additives designed to create special effects, such as colored flames, thick smoke, or sparks.
A hybrid rocket propulsion system typically comprises a solid fuel and a liquid or gas oxidizer. These systems are superior to solid propulsion systems in the respects of safety, throttling, restartability, and environmental cleanliness. However, hybrid systems are slightly more complex than solids, and consequently they are heavier and more expensive.
Common oxidizers include gaseous or liquid oxygen and nitrous oxide.
The Reaction Research Society (RRS), although known primarily for their work with liquid rocket propulsion, has a long history of research and development with hybrid rocket propulsion.
Several universities have recently experimented with hybrid rockets. BYU, the University of Utah and Utah State University launched a student-designed rocket called Unity IV in 1995 which burned the solid fuel Hydroxy-terminated polybutadiene (HTPB) with an oxidizer of gaseous oxygen, and in 2003 launched a larger version which burned HTPB with nitrous oxide.
Portland State University also launched several hybrid rockets in the early 2000's.
SpaceShipOne, the first private manned spacecraft, is powered by a hybrid rocket burning HTPB with nitrous oxide. The hybrid rocket engine was manufactured by SpaceDev. SpaceDev partially based its motors on experimental data collected from the testing of AMROC's (American Rocket Company) motors at NASA's Stennis Space Center's E1 test stand. Motors ranging from as small as 1000 lbf (4.5 kN) to as large as 250,000 lbf (1.1 MN) thrust were successfully tested. SpaceDev purchased AMROCs assets after the company was shut down due to lack of funding.
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nitrous oxide used in hybrids. Because just one propellant is a fluid, hybrids are simpler than liquid rockets.
Hybrid motors suffer two major drawbacks. The first, shared with solid rocket motors, is that the casing around the fuel grain must be built to withstand full combustion pressure and often extreme temperatures as well. Modern composite structures handle this problem well.
The primary remaining difficulty with hybrids is with mixing the propellants before burning. In solid propellants, the oxidizer and fuel are mixed in a factory in carefully controlled conditions (and even then it is tricky). Liquid propellants are generally mixed by the injector at the top of the combustion chamber, which directs many small fast-moving streams of fuel and oxidizer into one another. Liquid fuelled rocket injector design has been studied at great length and still resists reliable performance prediction. In a hybrid motor, the mixing happens at the surface of the melting or evaporating surface of the fuel. The mixing is not a well controlled process and generally quite a lot of propellant is left unburned, which limits the efficiency and thus the exhaust velocity of the motor.
There has been much less development of hybrid motors than solid and liquid motors. For military use, ease of handling and maintenance have driven the use of solid rockets. For orbital work, liquid fuels are enough better than hybrids that most development has concentrated there. There has recently been an increase in hybrid motor development for nonmilitary suborbital work:
•The Reaction
Research Society
(RRS), although
known primarily for
their work with
liquid rocket
propulsion, has a
long history of
research and
development with
hybrid rocket
propulsion.
•Several universities
have recently
experimented with
hybrid rockets.
BYU, the University
of Utah and Utah
State University
launched a student-
designed rocket
called Unity IV in
1995 which burned
the solid fuel
Hydroxy-terminated
polybutadiene
(HTPB) with an
oxidizer of gaseous
oxygen, and in 2003
launched a larger
version which
burned HTPB with
nitrous oxide.
•Portland State
University also
launched several
hybrid rockets in the
early 2000's.
•SpaceShipOne, the first private manned
spacecraft, is
powered by a hybrid
rocket burning
HTPB with nitrous
oxide. The hybrid
rocket engine was
manufactured by
SpaceDev. SpaceDev
partially based its
motors on
experimental data
collected from the
testing of AMROC's
(American Rocket
Company) motors at
NASA's Stennis
Space Center's E1
test stand. Motors
ranging from as
small as 1000 lbf
(4.4 kN) to as large
as 250,000 lbf (1.1
MN) thrust were
successfully tested.
The usual explanation for fuel-rich mixtures is that fuel-rich mixtures have lower
increases the ratio , which is approximately equal to the theoretical exhaust velocity. This explanation, though found in some textbooks, is wrong. Fuel-rich mixtures actually have lower theoretical exhaust velocities, because decreases as fast or faster than
ROCKET PROPELLANTS
•Introduction
•Liquids
•Solids
•Hybrids
•Tables of Properties
Propellant is the chemical mixture burned to produce thrust in rockets and consists
of a fuel and an oxidizer. A fuel is a substance which burns when combined with oxygen producing gas for propulsion. An oxidizer is an agent that releases oxygen for
combination with a fuel. Propellants are classified according to their state - liquid, solid, or hybrid.
The gauge for rating the efficiency of rocket propellants is specific impulse, stated in seconds. Specific impulse indicates how many pounds (or kilograms) of thrust are obtained by the consumption of one pound (or kilogram) of propellant in one second. Specific impulse is characteristic of the type of propellant, however, its exact value will vary to some extent with the operating conditions and design of the rocket engine.
Liquid Propellants
In a liquid propellant rocket, the fuel and oxidizer are stored in separate tanks, and are fed through a system of pipes, valves, and turbopumps to a combustion chamber where they are combined and burned to produce thrust. Liquid propellant engines are more complex then their solid propellant counterparts, however, they offer several advantages. By controlling the flow of propellant to the combustion chamber, the engine can be throttled, stopped, or restarted.
A good liquid propellant is one with a high specific impulse or, stated another way, one with a high speed of exhaust gas ejection. This implies a high combustion temperature and exhaust gases with small molecular weights. However, there is another important factor which must be taken into consideration: the density of the propellant. Using low density propellants means that larger storage tanks will be
required, thus increasing the mass of the launch vehicle. Storage temperature is also important. A propellant with a low storage temperature, i.e. a cryogenic, will require thermal insulation, thus further increasing the mass of the launcher. The toxicity of the propellant is likewise important. Safety hazards exist when handling, transporting, and storing highly toxic compounds. Also, some propellants are very corrosive, however, materials that are resistant to certain propellants have been identified for use in rocket construction.
Liquid propellants used by NASA and in commercial launch vehicles can be classified into three types: petroleum, cryogenics, and hypergolics.
Petroleum fuels are those refined from crude oil and are a mixture of complex hydrocarbons, i.e. organic compounds containing only carbon and hydrogen. The petroleum used as rocket fuel is kerosene, or a type of highly refined kerosene called RP-1 (refined petroleum). Petroleum fuels are used in combination with liquid oxygen as the oxidizer. Kerosene delivers a specific impulse considerably less than cryogenic fuels, but it is generally the best performer among the non-cryogenic options.
Liquid oxygen and RP-1 are used as the propellant in the first-stage boosters of the Atlas/Centaur and Delta launch vehicles. It also powered the first-stages of the Saturn 1B and Saturn V rockets.
Cryogenic propellants are liquefied gases stored at very low temperatures, namely liquid hydrogen (LH2) as the fuel and liquid oxygen (LO2 or LOX) as the oxidizer. LH2 remains liquid at temperatures of -253 degrees C (-423 degrees F) and LOX remains in a liquid state at temperatures of -183 degrees C (-298 degrees F).
Because of the low temperatures of cryogenic propellants, they are difficult to store over long periods of time. For this reason, they are less desirable for use in military rockets which must be kept launch ready for months at a time. Also, liquid hydrogen has a very low density (0.59 pounds per gallon) and, therefore, requires a storage volume many times greater than other fuels. Despite these drawbacks, the high efficiency of liquid oxygen/liquid hydrogen makes these problems worth coping with when reaction time and storability are not too critical. Liquid hydrogen delivers a specific impulse about 40% higher than other rocket fuels.
Liquid oxygen and liquid hydrogen are used as the propellant in the high efficiency main engines of the space shuttle. LOX/LH2 also powered the upper stages of the Saturn V and Saturn lB rockets as well as the second stage of the Atlas/Centaur launch vehicle, the United States' first LOX/LH2 rocket (1962).
Hypergolic propellants are fuels and oxidizers which ignite spontaneously on contact with each other and require no ignition source. The easy start and restart capability of hypergolics make them ideal for spacecraft maneuvering systems. Also, since hypergolics remain liquid at normal temperatures, they do not pose the storage problems of cryogenic propellants. Hypergolics are highly toxic and must be handled with extreme care.
Hypergolic fuels commonly include hydrazine, monomethyl hydrazine (MMH) and unsymmetrical dimethyl hydrazine (UDMH). The oxidizer is typically nitrogen tetroxide (N2O4 or NTO), though red-fuming nitric acid (RFNA) has also been used. RFNA has largely disappeared from use since the 1960s.
Despite the similarity of the names, hydrazine, MMH and UDMH are different compounds with unique chemical properties. Hydrazine gives the best performance as a rocket fuel, but it has a high freezing point and is too unstable for use as a coolant. MMH is more stable and gives the best performance when freezing point is an issue, such as spacecraft propulsion applications. UDMH has the highest freezing point and is stable enough to be used in large regeneratively cooled engines. Consequently, UDMH is often used in launch vehicle applications even though it is the least efficient of the hydrazine fuels. Also commonly used are blended fuels, such as Aerozine 50, which is a mixture of 50% UDMH and 50% hydrazine. Aerozine 50 is almost as stable as UDMH and provides better performance.
UDMH is used in many Russian, European, and Chinese rockets while MMH is used in the orbital maneuvering system (OMS) and reaction control system (RCS) of the Space Shuttle orbiter. The Titan family of launch vehicles and the second stage of the Delta II use Aerozine 50.
Hydrazine is also frequently used as a mono-propellant in catalytic decomposition engines . In these engines, a liquid fuel decomposes into hot gas in the presence of a catalyst. The decomposition of hydrazine produces temperatures of about 925 degrees C (1700 degrees F) and a specific impulse of about 230 or 240 seconds.
Solid Propellants
Solid propellant motors are the simplest of all rocket designs. They consist of a casing, usually steel, filled with a mixture of solid compounds (fuel and oxidizer) which burn at a rapid rate, expelling hot gases from a nozzle to produce thrust. When ignited, a solid propellant burns from the center out towards the sides of the casing. The shape of the center channel determines the rate and pattern of the burn, thus providing a means to control thrust. Unlike liquid propellant engines, solid propellant motors can not be shut down. Once ignited, they will burn until all the propellant is exhausted.
There are two families of solids propellants: homogeneous and composite. Both types are dense, stable at ordinary temperatures, and easily storable.
Homogeneous propellants are either simple base or double base. A simple base propellant consists of a single compound, usually nitrocellulose, which has both an oxidation capacity and a reduction capacity. Double base propellants usually consist of nitrocellulose and nitroglycerine, to which a plasticiser is added. Homogeneous propellants do not usually have specific impulses greater than about 210 seconds under normal conditions. Their main asset is that they do not produce traceable fumes and are, therefore, commonly used in tactical weapons. They are also often used to perform subsidiary functions such as jettisoning spent parts or separating one stage from another.
Modern composite propellants are heterogeneous powders (mixtures) which use a crystallized or finely ground mineral salt as an oxidizer, often ammonium perchlorate, which constitutes between 60% and 90% of the mass of the propellant. The fuel itself is aluminum. The propellant is held together by a polymeric binder, usually polyurethane or polybutadienes. Additional compounds are sometimes included, such as a catalyst to help increase the burning rate, or other agents to make the powder
easier to manufacture. The final product is rubberlike substance with the consistency of a hard rubber eraser.
Composite propellants are often identified by the type of polymeric binder used. The two most common binders are polybutadiene acrylic acid acrylonitrile (PBAN) and hydroxy-terminator polybutadiene (HTPB). PBAN formulations give a slightly higher specific impulse, density, and burn rate than equivalent formulations using HTPB. However, PBAN propellant is the more difficult to mix and process and requires an elevated curing temperature. HTPB binder is stronger and more flexible than PBAN binder. Both PBAN and HTPB formulations result in propellants that deliver excellent performance, have good mechanical properties, and offer potentially long burn times.
Solid propellant motors have a variety of uses. Small solids often power the final stage of a launch vehicle, or attach to payloads to boost them to higher orbits. Medium solids such as the Payload Assist Module (PAM) and the Inertial Upper Stage (IUS) provide the added boost to place satellites into geosynchronous orbit or on planetary trajectories.
The Titan, Delta, and Space Shuttle launch vehicles use strap-on solid propellant rockets to provide added thrust at liftoff. The Space Shuttle uses the largest solid rocket motors ever built and flown. Each booster contains 499,000 kg (1,100,000 pounds) of propellant and can produce up to 14,680,000 Newtons (3,300,000 pounds) of thrust.
Hybrid Propellants
Hybrid propellant engines represent an intermediate group between solid and liquid propellant engines. One of the substances is solid, usually the fuel, while the other, usually the oxidizer, is liquid. The liquid is injected into the solid, whose fuel reservoir also serves as the combustion chamber. The main advantage of such engines is that they have high performance, similar to that of solid propellants, but the combustion can be moderated, stopped, or even restarted. It is difficult to make use of this concept for vary large thrusts, and thus, hybrid propellant engines are rarely built.
PROPERTIES OF LIQUID ROCKET PROPELLANTS
Compound Chemical
Formula
Molecular
Weight
Density
Melting
Point
Boiling
Point
Liquid Oxygen O232.00 1.141 g/ml-218.8o C-183.0o C Hydrogen Peroxide H2O234.02 1.44 g/ml-0.4o C150.2o C Nitrogen Tetroxide N2O492.01 1.45 g/ml-9.3o C21.15o C Nitric Acid HNO363.01 1.55 g/ml-41.6o C83o C Liquid Hydrogen H2 2.0160.071 g/ml-259.3o C-252.9o C Dodecane (Kerosene)C12H26170.340.749 g/ml-9.6o C216.3o C Ethyl Alcohol C2H5OH46.070.789 g/ml-114.1o C78.2o C
Hydrazine N2H432.05 1.004 g/ml 1.4o C113.5o C Methyl Hydrazine CH3NHNH246.070.866 g/ml-52.4o C87.5o C Dimethyl Hydrazine(CH3)2NNH260.100.791 g/ml-58o C63.9o C NOTES:
Chemically, kerosene is a mixture of hydrocarbons; the chemical composition depends on its source, but it usually consists of about ten different hydrocarbons, each containing from 10 to 16 carbon atoms per molecule; the constituents include n-dodecane, alkyl benzenes, and naphthalene and its derivatives.
Nitrogen tetroxide and nitric acid are hypergolic with hydrazine, MMH and UDMH. Oxygen is not hypergolic with any commonly used fuel.
COMPOSITION OF SOLID ROCKET PROPELLANTS
Propellant Type Composition
Balistite (USA)
Double Base
Homogeneous
Nitrocellulose (51.5%), Nitroglycerine (43.0%), Plasticiser (1.0%), Other
(4.5%)
Cordite (Soviet)
Double Base
Homogeneous
Nitrocellulose (56.5%), Nitroglycerine (28.0%), Plasticiser (4.5%), Other
(11.0%)
SRB Propellant Composite
Ammonium Perchlorate (69.6%) as oxidizer, Aluminum Powder (16%) as
fuel, Iron Oxidizer Powder (0.4%) as catalyst, Polybutadiene Acrylic Acid Acrylonitrile (12.04%) as rubber-based binder, Epoxy Curing Agent
(1.96%)
NOTES:
The density of solid rocket propellants range from 1.5 to 1.85 g/ml. SRB propellant has a density。