Aeronautics and astronautics – Missile stabilization or trajectory control – Automatic guidance
Reexamination Certificate
1998-10-30
2001-10-30
Gregory, Bernarr E. (Department: 3662)
Aeronautics and astronautics
Missile stabilization or trajectory control
Automatic guidance
C244S003100, C244S003150, C244S003210
Reexamination Certificate
active
06308911
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates in general to the field of maneuver control of a vehicle traveling through a fluid environment and, more particularly, to a maneuver strategy to improve rapid turning capability.
2. Description of the Related Art
FIG. 1
illustrates a conventional vehicle
50
capable of traveling through a fluid medium, e.g., air, water, and plasma. The particular embodiment of the vehicle
50
illustrated is a missile system designed to travel through air. Although the invention is disclosed herein primarily in the context of a missile system, it is to be understood that the invention is not so limited. The invention is disclosed in the context of a missile system for the sake of clarity and to facilitate understanding of the invention. Those skilled in the art having the benefit of this disclosure will recognize that the teachings herein can be extrapolated to other vehicles capable of traveling through other fluid media, e.g., a torpedo traveling through a body of water.
The missile system
50
illustrated in
FIG. 1
can be described as an elongated body
100
that travels through a fluid medium, typically the earth's atmosphere. The missile
100
has a forward section
102
and an aft section
104
divided by the missile body
100
's center-of-gravity
105
, the forward section
102
having a control device
110
. Missile control devices can fundamentally be categorized as forward aerodynamic, e.g., canards; forward propulsive, e.g., a thruster; aft aerodynamic, e.g., fins; and aft propulsive, e.g., thrust vector control. Thus, it will be apparent to those skilled in the art having the benefit of this disclosure that other control devices alternative to those shown in
FIG. 1
are possible and may even be desirable depending on the particular application for the missile system
50
. For instance, the forward control device
110
, a set of thrusters in the embodiment illustrated, may be implemented with canards, instead. Control devices may also be used in tandem.
Certain aspects of a missile's maneuverability are typically expressed in terms of angles.
FIG. 2
illustrates the angle nomenclature typically used by those in the art to describe missile maneuverability. The nomenclature is frequently defined in terms of an inertial coordinate system that is fixed relative to the earth. An axis in this coordinate system is an “inertial axis.” For simplicity, a single pitch plane will be used to define the nomenclature although those in the art will recognize that more than one plane might be relevant in some circumstances. The angle
200
between the nose
215
of the missile system
50
and the inertial axis
205
is called the pitch angle. The angle
210
between the nose
215
of the missile
50
and the missile's velocity vector
220
is called the angle-of-attack. The angle
225
between the velocity vector
220
, i.e., the magnitude and direction of the missile's travel, and the inertial axis
205
is called the flight path angle. Thus, as shown in
FIG. 2
, for pitch plane maneuvers, the pitch angle
200
is the sum of the angle-of-attack
210
and the flight path angle
225
.
Missiles are typically directed toward a target. This target may be in the form of a set of inertial coordinates or a physical object or a combination of both. If the missile is stationary at launch, its initial flight path angle
225
, is aligned with the missile's nose or equal to its pitch angle
200
. If the missile is air launched, the missile's initial flight path angle is inherited by the flight path angle of the host aircraft and launcher imposed dynamics. The initial flight path angle
225
is typically not equal to the flight path angle desired to reach the target. The missile must maneuver to change its current flight path angle
225
to its desired flight path angle.
Typically, a missile system relies on aerodynamic force to maneuver. In the atmosphere, the amount of force the missile system can receive from aerodynamics exceeds the amount of lateral control force that can be packaged in the missile at altitudes of less than about 20 km. Because of this, the primary purpose of control mechanisms is generating a moment to rotate the nose of the missile to produce the angle-of-attack
210
. The angle-of-attack
210
, in turn, generates an acceleration normal to the missile body which results in a rotation of the missile's flight path angle
225
. Maintaining an adequate missile velocity to supply the necessary aerodynamic forces and moments is required to maneuver in the conventional manner.
More particularly, and referring again to
FIG. 1
, a force generated by a control device, such as the forward control device
110
, generates a moment, or torque, on the missile body
100
about the center-of-gravity
105
of the missile body
100
. Below an altitude of approximately 20 km, a missile system's primary source of transverse acceleration is the aerodynamic force resulting from the missile body
100
's angle-of-attack. Flight control devices, e.g., forward thrusters
110
, are commanded by control processes encoded in and executed by computational devices carried by the missile system to generate moments and forces and obtain this angle-of-attack.
FIG. 3
is a functional block diagram of a missile control system
301
for the conventional missile system
50
of FIG.
1
. Block
330
represents the physical system, i.e., the missile, and incorporates descriptions of all vehicle subsystems including, for example, control actuation, propulsion, and inertial measurement systems as well as the aerodynamic configuration. The missile guidance logic
310
determines a desired kinematic trajectory by commanding a dynamic response that results in a change to the missile's flight path. The vehicle's measured dynamic response is shown as feedback signal
305
. This signal encodes, for example, a measurement of the missile's rotational rates, translational rates, accelerations, and inertial orientation. In general terms, the autopilot controller
325
is a control process that uses the difference between desired missile responses, e.g., rates, accelerations, or attitudes, and the measured responses to define a set of error signals
320
which the controller uses to encode control commands
330
for a control actuation device, e.g., thruster level, that is a part of the vehicle dynamics and kinematics
300
.
The energy necessary to operate a flight control device is very precious in an operational sense. A missile system can carry only a limited amount of energy, e.g., energy stored as rocket fuel, batteries, etc. Strategically, altering the amount of energy that may be provided can affect the missile system's range and performance against a target. Tactically, the energy is conserved as much as possible so that the missile system can retain its ability to maneuver from the time it is deployed to the time it reaches the target. Thus, the energy necessary to operate the flight control device is an important design consideration for both the physical structure of the missile system and its operation.
Air launched missiles are typically mounted in a forward facing direction where the nose of the missile is aligned with the nose of the host aircraft. A missile able to reverse the direction of its flight path, e.g., 180° turn, is able to provide a capability to destroy rearward approaching aircraft before they are able to be within firing range of the host aircraft. A combat advantage is also provided in a fly-by scenario where an enemy aircraft passes the host aircraft traveling the opposite direction. If the missile can turn rapidly enough, the enemy aircraft can not escape being defeated. Thus, the time which the maneuver consumes, the velocity of the missile after the maneuver, and the turning radius or distance which the missile deviated from its desired flight path is critical to the life and death nature of the engagement.
Thus, one application exhibiting an immediate n
Gregory Bernarr E.
Lockheed Martin Corp.
Sadacca Stephen S.
Sidley Austin Brown & Wood
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