Aeronautics and astronautics – Spacecraft – With fuel system details
Reexamination Certificate
2000-08-09
2001-12-25
Poon, Peter M. (Department: 3643)
Aeronautics and astronautics
Spacecraft
With fuel system details
C244S003230, C244S164000, C102S374000
Reexamination Certificate
active
06332592
ABSTRACT:
FIELD OF THE INVENTION
The present invention concerns velocity precision pointing in spin-stabilized spacecraft during thrusting maneuvers, which overcomes imperfections, such as small misalignments and offsets, in the spacecraft introduced during manufacture or operation thereof which cause velocity pointing error.
BACKGROUND OF THE INVENTION
Ideally, during an axial thrust maneuver, an axial force is applied by the thrusters through the center of the mass to propel the spacecraft in a desired direction. However, as shown in
FIG. 1
, when a spinning rocket performs an axial thrusting maneuver, there are always body-fixed torques due to various error sources such as thruster misalignment and center-of-mass offset. The angular momentum vector remains inertially fixed, unless acted upon by an external torque.
Because of these body-fixed torques, the angular momentum vector traces a circular path in inertial space, i.e., an internal wobble is produced in the spacecraft. This internal wobble, in turn, produces an error in the pointing of the spacecraft, known as the velocity pointing error.
In spacecraft and rocket dynamics and control, two main methods are commonly used to point the vehicle along the desired path (trajectory) and obtain stability of the vehicle, i.e., reduce pointing error. When the vehicle travels in exactly the right trajectory, there is said to be no pointing error.
One method to obtain stability and an acceptable trajectory is referred to as 3 axis stabilized control, in which the spacecraft or rocket is not spinning or is spinning at an extremely low rate. In this method, a feedback control loop is provided which, while the engine is firing, sensors detect the direction in which the vehicle is actually pointing, and compares the actual direction with the intended direction.
To correct pointing errors, small thrusters are fired or the engine is rotated on a gimbal to counteract unfavorable forces. However, to implement these methods, additional thrusters must be installed on the vehicle or a complicated gimbal mechanism must be built into the spacecraft or rocket. Either of these methods add considerable extra weight, complexity, increase the chances for malfunction and increase cost of construction, launch and operation.
In the second method, spacecraft and rockets are often spun to provide stability and obtain a correct trajectory. Most commonly, spin rates are increased to overcome the error. In a representative example, the Galileo spacecraft, before performing an axial maneuver, increased the spin of the spacecraft from 3 rpm's to 10 rpm's to provide greater stability.
However, to reduce the velocity-pointing error of spacecraft or rockets, either considerable additional fuel must be carried into space to increase the spin rate, or improvements in tolerances during the manufacture of the spacecraft engines are needed, both of which are very costly.
In another proposed method, as disclosed in the article entitled “Annihilation of Angular Momentum Bias During Thrusting and Spinning-up Maneuvers”,
The Journal of the Astronautical Sciences
, vol. 37, No. 4, October-November 1989, pp. 433-450, a maneuver scheme involving two burns, with a coast or delay in between, may be used. After the first burn, when the thruster reaches the appropriate orientation, the spacecraft coasts. Then, the thruster is reignited, initiating a second burn, which causes the angular momentum vector to encircle the inertial Z axis. However, this maneuver bars the use of a solid fuel motor.
In view of the above, it is an object of the present invention to overcome the above mentioned problems/disadvantages encountered when attempting to overcome velocity pointing error in spin-stabilized vehicles, such as spacecraft or rockets, during thrusting maneuvers.
It is another object of the present invention to produce a thrust profile to minimize velocity pointing error while simultaneously driving the angular momentum of the spacecraft or rocket back to its originally desired position.
It is a further object of the present invention to eliminate the need for high spin rates, which require additional fuel for producing high spin rates and increases cost, while obtaining velocity precision pointing with a pointing error up to several magnitudes better than that of the presently used methods.
SUMMARY OF THE INVENTION
In this regard, the present inventors earnestly conducted research in order to overcome the disadvantages of the above methods of reducing velocity pointing error in spin-stabilized spacecraft, and unexpectedly discovered that by altering a spacecrafts' thrust profile, the ramp up time, i.e., the time needed to gradually increase thrust over a time period t
r
to rise from a low value of from 0 to 50% of maximum thrust to a high value of greater than 50% of maximum thrust, by softening the ignition transient, produces an astounding and unexpected reduction in velocity pointing error. Specifically, the ramp up time must be equal to or greater than one period of revolution of the spacecraft.
To achieve such an object, in a first embodiment of the present invention, there is disclosed a method for velocity precision pointing in spin-stabilized spacecraft or rockets during thrusting maneuvers comprising gradually increasing thrust over a time period t
r
to rise from a low value of from 0 to 50% of maximum thrust to a high value of greater than 50% of maximum thrust by softening the ignition transient, the time period of the rise in thrust t
r
being equal to or greater than at least one period of the spin rate of the spacecraft, and being calculated by the following formula (I):
t
r
=n(2&pgr;/&OHgr;) (I)
wherein n, the number of revolutions of the spacecraft desired, is a number greater than or equal to 1 and less than or equal to t
b
&OHgr;/4&pgr;, t
b
being the total burn time and &OHgr;, the spacecraft spin rate, is greater than 0 and less than or equal to 42 radians/second.
In another embodiment of the present invention based on the first embodiment above, in spin-stabilized spacecraft or rockets having solid rocket motors using solid propellant, softening of the ignition transient is achieved by modification of the grain geometry of the solid propellant to delay burn time so that t
r
=n(2&pgr;/&OHgr;).
In another embodiment of the present invention based on the first embodiment above, in spin-stabilized spacecraft or rockets having a solid rocket motor using solid propellant, softening of the ignition transient is achieved by modification of the solid propellant comprising applying a coating of slow burning material upon the solid propellant. Such slow burning material may include waxes, binders, or any material to slow the burn rate of the propellant.
In another embodiment of the present invention based on the first embodiment above, in spin-stabilized spacecraft or rockets having a solid rocket motor using solid propellant having an exposed surface area, softening of the ignition transient is achieved by varying the grain density, the density of fuel and oxidizer in a binder, of the solid propellant over an exposed surface area or as a function of propellant depth, where depth is defined in the direction of the burn surface area regression.
In another embodiment of the present invention based on the first embodiment above, in spin-stabilized spacecraft or rockets having a combustion chamber, the combustion chamber is pre-pressurized with a gas having a molecular mass smaller than that of the combustion products produced in the combustion chamber, and which is also unreactive with the propellant to be used.
In another embodiment of the present invention based on the first embodiment above, in spin-stabilized spacecraft or rockets having a combustion chamber, one or more relief valves is installed in the combustion chamber to soften the ignition transient by regulating pressure in the combustion chamber.
In another embodiment of the present invention based on the first embodiment above, in spin-stabilized spacecraft or rockets having a
Javorsek, II Daniel
Longuski James M.
Collins Timothy D.
Poon Peter M.
Townsend Donald E.
Townsend & Banta
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