Power plants – Reaction motor – Method of operation
Reexamination Certificate
2000-05-24
2002-03-12
Thorpe, Timothy S. (Department: 3746)
Power plants
Reaction motor
Method of operation
C060S231000, C060S251000
Reexamination Certificate
active
06354074
ABSTRACT:
FIELD OF THE INVENTION
This invention relates to control of pitch and yaw of vehicles propelled by high-velocity gas jets, as for example rockets.
BACKGROUND OF THE INVENTION
Rocket and other jet-type propulsion systems are widely used for vehicle propulsion, notably for lifting payloads into orbit, for destructive military missiles, and for military aircraft. In general, such systems or engines produce thrust by discharge of a plume or exhaust at high velocity along the axis of a nozzle. The problems of pitch and yaw directional control of such vehicles are widely known, and have been solved in a number of ways. External fins and canards have been used for directional control. These fins may be fixed for the most general directional control, or they may include articulable flaps which are controlled in response to a controller for feedback flight control. When the vehicle equipped with fins or canards must itself be carried on an aircraft before launch, or mounted in a canister, the fins and or canards may be arranged in a stowed configuration, and deployed in conjunction with the initial stages of launch. Such fins and canards depend upon aerodynamic forces, so are only usable within the atmosphere, which may be taken to extend to an altitude of 100,000 feet. Within the atmosphere, the use of fins or canards increases the drag of the vehicle equipped therewith.
Another form of attitude or directional control of a vehicle equipped with rocket-type propulsion is that of thrust vector control (TVC), described U.S. Pat. No. 2,943,821, issued Jul. 5, 1960 in the name of Wetherbee, Jr.; U.S. Pat. No. 3,166,897, issued Aug. 21, 1961 in the name of Lawrence et al; in U.S. Pat. No. 3,132,476, issued May 12, 1964 in the name of Conrad; and U.S. Pat. No. 3,132,478, issued May 12, 1964 in the name of Thielman, and in the text “Rocket Propulsion Elements” by Robert Sutton. Sutton categorizes TVC mechanisms into four basic categories:
(a) Mechanical deflection of a nozzle or thrust chamber;
(b) insertion or adjustment of vanes located in the jet exhaust stream;
(c) injection of fluid into the diverging nozzle section to deflect the exhaust flow; and
(d) separate thrust-producing devices which are independent of the main flow through the nozzle, providing two thrust vectors which may be summed to obtain the net thrust vector.
It should be noted that this last may not be a form of TVC, since it does not act on the thrust vector itself, but merely adds a separate thrust vector.
Mechanical deflection of a nozzle or thrust chamber requires a highly reliable movable structure which is subject to the entire thrust load, which may be costly and undesirably massive. Insertion or adjustment of vanes within the exhaust stream requires vanes which are structurally sound at the very high temperatures and pressures of the exhaust stream. Thrust-producing devices independent of the main nozzle have been used, especially for end-of-flight corrections of destructive missiles acting against maneuvering targets; they must, however, be located within the body of the vehicle if additional aerodynamic drag is to be avoided.
According to Sutton, “the injection of secondary fluid through the wall of the nozzle into the main gas stream has the effect of forming oblique shocks in the nozzle diverging section, thus causing a deflection of part of a main gas flow,” and this deflection of the main gas flow, in turn, results in a deviation of the thrust vector from the axis of the nozzle.
Liquid injection thrust vector control is described in U.S. Pat. No. 3,737,103, issued Jun. 5, 1973 in the name of Howell et al. Liquid injection thrust vector control is a proven technology, which is used in applications such as Titan III and Minuteman. In liquid injection thrust vector control, liquid is stored in either the propellant tanks or auxiliary tanks of the vehicle. The liquid is controllably distributed or manifolded to various injection positions around the periphery of the nozzle. When a pitch or yaw correction is desired, a signal is sent to a valve or valves controlling the injection of liquid into the exhaust plume at locations associated with the plane(s) of the correction thrust. Injection of the liquid into the exhaust stream results in vaporization of the liquid, and also results in a change in thrust along the relevant plane. Liquid injection has known problems, which include the instability of stored liquids, as described in U.S. Pat. No. 3,092,963, issued Jun. 11, 1963 in the name of Lawrence. Also, the axial thrust of the vehicle is reduced by the energy required to vaporize the injected liquid, and to bring it up to the temperature of the surrounding gas. The amount of liquid which is required to produce a given change in attitude is generally determined by experimentation.
FIG. 1
is a chart illustrating the amount of side injectant which is required to produce a side force, according to Sutton. In
FIG. 1
, the ordinate- or y-axis represents the ratio of the side force divided by (or normalized to) axial force, and the abscissa- or x-axis represents the ratio of injectant mass flow divided by primary mass flow. As illustrated in
FIG. 1
, injection of inert liquids results in the least side thrust or directional control for a given mass flow, while reactive fluids provide greater control.
The greatest control is provided by a flow of propellant hot gas. Such charts can be obtained experimentally by maintaining a constant main exhaust flow rate through the nozzle, while varying the side injection flow rate.
As suggested by
FIG. 1
, greater side force or thrust control can be achieved by injection of reactive liquids than of inert liquids. U.S. Pat. No. 2,952,123, issued Sep. 3, 1960 in the name of Rich, describes injection of fuel into a jet nozzle, which burns in the supersonic exhaust stream to provide directional control. Similarly,
FIG. 1
indicates that injection of propellant hot gas provides yet greater side force or directional control as a function of mass flow.
As an example of the use of the chart of
FIG. 1
, consider a rocket engine or motor which generates an axial thrust of 10,000 pounds force (lbf) at an exhaust flow rate of 33 pounds of mass (lbm). If a 2° deflection of the thrust vector is required, then (10,000) (sin 2°) of lateral force, corresponding to 349 lbf of normal side force, is required. The side-to-axial-force ratio is calculated as 0.035, which corresponds to a side injection flow rate of (0.035×33)=1.98 lbm/s for inert liquids, 1.32 lbm/s for reactive liquids, or 1.0 lbm/s for propellant hot gas.
From
FIG. 1
, it is apparent that injection of hot gas is the most effective way, in terms of relative mass flow, to achieve side force or directional control. One advantageous way to provide hot gas for side injection is to tap the gas from the main combustion chamber, because the chamber pressure is greater than the static pressure in the nozzle as a result of expansion, and a substantial side injection flow rate can therefore be achieved. U.S. Pat. No. 3,759,039, issued Sep. 18, 1973 in the name of Williams, describes the bleeding of hot gases from the combustion chamber of a rocket, by way of controllable valves, into the side of the nozzle. In such an arrangement, the valves must control the flow of very hot gases, which may adversely affect their reliability, and may result in a costly structure using exotic materials and sacrificial elements or coatings.
The problem of control of the flow of very hot propellant-type gases makes the use of cooled or cold gas advantageous. The chamber temperatures of liquid- and solid-propellant rockets may approach 6000° F., which is too high for conventional piping and valves. Cool-gas or cooled-gas injection is described in U.S. Pat. No. 3,255,971, issued Jun. 14, 1966 in the name of Widell; U.S. Pat. No. 3,698,642, issued Oct. 17, 1972, in the name of McCullough; U.S. Pat. No. 4,384,694 in the name of Watanabe et al.; U.S. Pat. No. 4,424,670, issued Jan. 10, 1984 in the name of Calabro; and in the abovementioned Lawrence pa
Arves Joseph Paul
Jones Herbert Stephen
Kearney Darren Andrew
McLeod Rory Nell
Roberts Ryan Earl
Fierke T. G.
Gartenberg Ehud
Lockheed Martin Corp.
Meise W. H.
Thorpe Timothy S.
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