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PREFACE
CHAPTER 1
CLASSIFICATION
DUCT JET PROPULSION
The uniqueness of the missile, e.g. high thrust to weight, high thrust head-on. Ramjets depend on rocket boosters or some other method (such as being launched from an aircraft) to be accelerated to near their design airspeed in order to become functional.
ROCKET PROPULSION
However, other terms, such as pusher cylinder and combustor, are still used in the literature. An example is a pressurized liquid propellant system that uses a solid propellant to generate hot gases to pressurize the reservoir; Flexible diaphragms are needed to separate the hot gas and liquid jet propellant in the tank.
APPLICATIONS OF ROCKET PROPULSION
Details of several of these Space Shuttle rocket propulsion systems are given elsewhere in this book. Ziolkowsky, Space Investigations by Means of Propulsive Spaceships (in Russian), Kaluga, Saint Petersburg, 1914. Eds.), The Papers of Robert H.
CHAPTER 2
DEFINITIONS AND FUNDAMENTALS
- DEFINITIONS
- THRUST
- EXHAUST VELOCITY
- ENERGY AND EFFICIENCIES
- TYPICAL PERFORMANCE VALUES
The first term is the thrust represented by the product of the mass flow of the propellant and its exhaust velocity relative to the vehicle. When p2 = p3 the effective exhaust velocity c is equal to the average true exhaust velocity of the propellants v2.
CHAPTER 3
NOZZLE THEORY AND
THERMODYNAMIC RELATIONS
IDEAL ROCKET
These equations theoretically describe a quasi-one-dimensional nozzle flow, which corresponds to an idealization and simplification of the full two- or three-dimensional equations and the real aerothermochemical behavior. Later in this book we present more sophisticated theories or introduce correction factors for some of the items in the list and they allow a more precise determination of the simplified analysis.
SUMMARY OF THERMODYNAMIC RELATIONS
In the English Engineering System of Units, a different constant (see Appendix 1) must be specified to account for the units of mass (ie lbm). The relationship between the stagnation pressure and the local pressure in the flow can be found from the two previous equations:
ISENTROPIC FLOW THROUGH NOZZLES
This equation shows that the thrust is proportional to the throat area At and the chamber pressure (or nozzle inlet pressure) PI and is a function of the pressure ratio across the nozzle piJp2, the specific heat ratio k and the pressure thrust. The nozzle would not flow fully below an area ratio of about 6 or 7 and the gas jet would only be in the center of the exit area. Correct the thrust value for sea level rise and the corresponding specific impulse.
NOZZLE CONFIGURATIONS
However, the contour of the nozzle wall is different and the change in cross-sectional area per unit length decreases. The middle set of curves shows the starting angle 8; and exit angle ee as a function of nozzle area ratio and percent length. As the ambient pressure decreases, the hot gas stream fills an increasing portion of the diverging part of the nozzles.
REAL NOZZLES
This causes a decrease in the gas flow per unit area and the transfer of the latent heat of vaporization to the remaining gas. The velocity correction factor Sv is defined as the square root of the energy conversion efficiency y'e. This factor is also approximately the ratio of the actual specific impulse to the ideal or theoretical specific impulse.
FOUR PERFORMANCE PARAMETERS
Since the specific impulse is proportional to the exhaust velocity, its actual value can be found by multiplying the theoretical value by the velocity correction factor t;V'. They usually refer to conditions that allow immediate evaluation or comparison with reference values, and often refer to conditions that can be easily measured and/or corrected. This minimum value can be determined by probabilistic evaluation of these losses and then usually confirmed by actual static and in-flight tests.
NOZZLE ALIGNMENT
The determination of this value may be based on a nominal value (items I or 2 above) reduced by all possible losses, including changes in chamber pressure due to changing pressure drop in the injector or piping, a loss due to nozzle surface roughness, propellant initial ambient temperatures, production variations from rocket to rocket (eg, in grain volume, nozzle dimensions or pump impeller diameters, etc.). For simple unguided rocket vehicles, it has been customary to roll or rotate the vehicle to prevent deflection from being in one direction only or to mitigate distortion during powered flight. In this case, the nozzles are cut at an angle to the surface of the vehicle, which allows for a compact installation.
VARIABLE THRUST
For an ideal rocket with a characteristic velocity c* = 1500 m/sec, a nozzle throat diameter of 18 cm, a thrust coefficient of 1.38, and a mass flow rate of 40 kg/sec, calculate the chamber pressure, the push, and the specific impulse. If this chamber pressure is doubled, what happens to the thrust and the exit velocity. For an ideal rocket with a characteristic velocity c* of 1220 m/sec, a mass flow rate of 73.0 kg/sec, a thrust coefficient of 1.50, and a nozzle throat area of 0.0248 m2, calculate the effective exhaust velocity, the thrust, the chamber pressure, and the specific impulse.
FLIGHT PERFORMANCE
GRAVITY-FREE, DRAG-FREE SPACE FLIGHT
For a rocket where the propellant flow rate is constant, the instantaneous mass of the vehicle m can be expressed as a function of the initial mass of the complete vehicle m0, mp, tp and the instantaneous time t. The concept of the maximum attainable flight speed increase /";.u in a zero-gravity vacuum is useful for understanding the influence of the basic parameters. A mass fraction of 0.80 would indicate that only 20% of the vehicle's total mass is available for structure, skin, cargo, propulsion, radios, steering system, aerodynamic lift surfaces, and so on; the remaining 80% is useful propellant.
For low flight speeds, the effect of Mach number can be neglected and the drag and lift coefficients are functions of the angle of attack. The variation of drag and lift coefficients for a typical supersonic rocket is shown in Fig. The forces of gravity pull the vehicle towards the center of mass of the towing body.
BASIC RELATIONS OF MOTION
Let e be the angle of the flight path with the horizontal and ijJ the angle of the pressure direction with the horizontal. 4--5 show the net force (by adding thrust, drag, and gravity vectors) that is at an angle to the flight path, which will be curved. Assume that the local gravitational acceleration is equal to sea level g0 and is invariant throughout the flight.
EFFECT OF PROPULSION SYSTEM ON VEHICLE PERFORMANCE
One way is by reducing the final mass m1, which consists of the inert hardware plus the unusable, residual mass of the propellant. It is influenced by the design of the nozzle exit (exit pressure) and the geometry of the vehicle base model. The length of the thrust nozzle is often an important part of the overall length of the vehicle or stage.
SPACE FLIGHT
7; here the Earth (or any celestial body around which another body moves) is at one of the foci of this ellipse. Satellites at medium and low altitudes (500 to 35,000 km) experience disturbances due to the flattening of the Earth. Depending on the inclination of the orbital plane relative to the Earth's equator and the height of the satellite's orbit, two perturbations arise: (1) the regression of the nodes, and (2) displacement of the apside line (major axis).
FLIGHT MANEUVERS
In translational maneuvers, the rocket propulsion force vector passes through the vehicle's center of gravity. It can be performed by the main propulsion system in the upper stage of the launcher. An RCS can be incorporated into the payload stage and each of the stages of a multi-stage vehicle.
FLIGHT VEHICLES
Determine the payload for two cases: (1) when the masses of the two stages are equal, and (2) when the mass ratios of the two stages are equal. When the mass ratio of the stages is equal, the payload is a maximum for gravity-free vacuum flight and the distribution of the masses between. The thrust sizes depend on the mass of the vehicle, which in turn depends on the mass of the payload and the mission.
MILITARY MISSILES
For more accurate values, the velocity increase u0 is the initial velocity of the launching aircraft. In a particular air-to-air combat situation, the effectiveness of the rocket projectile varied approximately inversely as the cube of the time spent on target. The analysis of the missile and propulsion configuration that provides the minimum time to target over all the likely flight scenarios can be complex.
AERODYNAMIC EFFECT OF EXHAUST PLUMES
It can be a single rocket propulsion system that has a short high initial thrust and a smaller (10 to 25%) sustaining thrust of lower duration. Drag losses can be reduced if the missile has a large L/D ratio (or a small cross-sectional area) and if the propellant density is high, allowing a smaller missile volume. The drag forces can be high if the missile is traveling at low altitude and high speed.
FLIGHT ST ABILITY
It is possible to exercise control over the movement of the center of gravity by well-considered design. The designer generally has less freedom to control the movement of the center of gravity for solid propellant rockets. What is the mass ratio mp/mo for a vehicle that has one-fifth of its original initial mass at the time of the end of the rocket operation.
CHAPTER 5
CHEMICAL ROCKET PROPELLANT PERFORMANCE ANALYSIS
BACKGROUND AND FUNDAMENTALS
The specific heat ratio k of the mixture can be determined from a similar summary or from Eq. In general, the heat of reaction can be determined from the sums of the heats of formation of the products and reactants, respectively. G for reactions at constant temperature and pressure is the chemical potential of the products less that of the reactants.
ANALYSIS OF CHAMBER OR MOTOR CASE CONDITIONS
However, three equations solve only three unknowns, for example the combustion temperature and the molar fractions of two of the species. If they are not in equilibrium, another combustion temperature value is chosen until there is convergence and the energy is in equilibrium. For solid propellants, combustion efficiency is a function of the pellet design, the propellant and the degree of mixing between the various solid components.
ANALYSIS OF NOZZLE EXPANSION PROCESSES
Some propellant products include species that condense as the temperature drops as the nozzle expands. This loss of chamber pressure results in a slight decrease in c and ls values. The change in gas enthalpy of the hot gas in the combustion chamber is numerically equal to the heat of formation.
COMPUTER ANALYSIS
Time-dependent chemical reactions in the chamber are usually neglected, but they can be analyzed by estimating the time rate at which the reaction occurs; one way is to calculate the time derivative of the advance rate d}.,/dt and then set this derivative to zero. The main assumptions for this program are the one-dimensional forms of the continuity, energy, and momentum equations, zero velocity at the leading edge of the chamber, isentropic expansion at the nozzle, use of the ideal gas laws, and chemical equilibrium in combustion. room. The large pressure drop in the chamber (approximately 126 psi) is due to the energy required to accelerate the gas, as discussed in Section 3.3 and Table 3-2.
RESULTS OF THERMOCHEMICAL CALCULATIONS
Aggressive gases, such as 02, 0 or OH, can cause oxidation of the wall materials of the chamber and nozzle. The influence of the degree of expansion or nozzle exit pressure on the gas composition is shown in fig. Dissociation of the reaction products increases as the chamber temperature increases and decreases with increasing chamber pressure.
