Aeronautics and astronautics – Composite aircraft
Reexamination Certificate
2002-01-29
2003-01-21
Swiatek, Robert P. (Department: 3644)
Aeronautics and astronautics
Composite aircraft
C244S063000, C244S172200
Reexamination Certificate
active
06508435
ABSTRACT:
FIELD OF THE INVENTION
The invention relates to aerospace technology, in particular, to methods for orbital injection of aerospace systems to put different payloads, such as communication, navigation, and monitoring satellites (including ecology monitoring satellites) into low and medium earth orbits, as well as to promptly deliver payloads to remote terrestrial and oceanic areas.
BACKGROUND OF THE INVENTION
A method is known for orbital injection of payloads (PL) by an aerospace system (ASS) comprising a carrier aircraft (CA), a mid-stage booster, and an orbiter as a payload.
The method presumes the horizontal flight of a CA, its climb to an altitude of about 20 km and acceleration to a flight speed of 800-1100 km/h. After this speed is reached, the motors of the mid-stage booster are ignited and it is separated from the CA. Then the mid-stage booster accelerates and delivers the PL to the targeted orbital injection point, after which the PL separates from the booster (RU Pat. No. 2061630, IPC B64G {fraction (1/14)} .
A drawback of this method is the risk of the mid-stage booster firing prior to its separation from the CA, as well as the impossibility to reach the whole system's potential weight-lifting capability because the mid-stage booster is fired in the CA horizontal flight mode.
A method is known for orbital injection, including a horizontal takeoff system comprising a tanker airframe with a powered separable aircraft (A/C).
This method is realized in the following manner. The horizontal takeoff of the whole system is provided by the A/C engines supplied with fuel from airframe tanks. After attaining the predetermined operating speed, the A/C separates from the airframe and enters the prescribed trajectory, and the airframe returns to Earth (RU Pat. No. 2120398, IPC
6
B64G {fraction (1/14)}, applicant—DASA, a German company).
This method has the same disadvantages as the method described above, along with poor energetic efficiency due to the necessity of using the A/C sustainers to boost the entire airframe—A/C system off the ground, which results in an unjustified overweight of the A/C structure, and a loss in its load-carrying capacity.
A method is also known for putting an aerospace system, Rockwell International Corporation, into orbit (U.S. Pat. No. 5,402,965IPC
6
B64G {fraction (1/14)}). In accordance with this method, a carrier aircraft with a launch vehicle (LV), consisting of a winged first stage mated with a payload and a recoverable winged last stage, carries out horizontal flight to the LV launching point, the LV separates from the CA, the LV is fired into the scheduled trajectory, and the recoverable winged last stage is separated from the LV. The recoverable winged last stage inserts into orbit, performs the mission task, executes de-orbitation, atmospheric aerodynamic deceleration, and controlled aircraft-like landing at a designated aerodrome.
A drawback of this method is the necessity of employing a winged first stage of the LV and separating it from the CA in horizontal flight, which correspondingly increases the structure weight of the first stage of the LV and does not make it possible to realize the optimum initial kinematic launch parameters of the LV after the LV-CA separation.
The analog most similar to the proposed method is the method for orbital injection disclosed in U.S. Pat. No. 4,901,949, IPC
6
B64C 3/38, Orbital Sciences Corporation (PCT/US 89/00867, Mar. 8, 1989) (protected by RU Pat. No. 2026798, IPC
6
B64D 5/100, F42B 15/00).
This method is used for an aerospace system comprising a CA, a three-stage LV with a winged first stage, and a PL.
The method consists in that an LV is accelerated to the point of its launch in the flight trajectory of the CA, the direction of flight of the CA coinciding with the direction of launching the LV, horizontal separation of the LV from the CA takes place, the motor system of the first stage is fired after separation and after the LV is lagging behind the CA in a horizontal position, the LV is boosted with use of the aerodynamic lift force of the winged first stage and the thrust force of its motors, after which the winged first stage is separated, and the second stage if fired.
One of the drawbacks of this method of orbital injection is the necessity of employing a wing in the first stage of the LV, which increases the weight of its construction and complicates the simultaneous control of aerodynamic and reactive actuating devices. Ignition of the motor system of the first stage when the LV is in a horizontal position prevents the CA from executing, prior to the LV-CA separation, an optimal maneuver to achieve design kinematic motion parameters (altitude, speed, trajectory pitch angle) ensuring the maximum lifting capability of the LV in the predetermined launch trajectory point.
DISCLOSURE OF THE INVENTION
The object of the present invention is to enhance an aerospace system's lifting capacity when putting a payload into space and delivering it to any terrestrial and oceanic areas, to ensure reliable LV-CA separation, to provide safety for the plane and crew members when firing the sustainers of the first stage of the LV, and to reduce the cost of launching a payload.
This object is accomplished in a method for controlling an aerospace system to put a payload into an orbit, comprising starting a carrier aircraft with a launch vehicle and a payload on board from a base aerodrome, flying it to a launch vehicle launch area, separating the launch vehicle from and lagging it behind the carrier aircraft with subsequent launch of the launch vehicle to a predetermined trajectory point and separation of the payload from the launch vehicle, the carrier aircraft at the maximum cruising speed mode in the launch area of the launch vehicle performs a dive to gain a maximum permissible horizontal flight speed, at the moment that speed is reached by the carrier aircraft, pitchup is effected at the maximum possible angle of attack, terminating with transition to an angle of attack ensuring a near-zero normal g-loading, wherein the aforesaid pitchup parameters are chosen so as to be adequate to the achievement, at a preset time t
p
, of a flight trajectory point where the design flight speed V
D
, flight altitude H
D
and trajectory pitch angle &phgr;
D
provide the maximum launch vehicle payload, as well as to ensure subsequent flight of the carrier aircraft with allowable parameters both after separation of the launch vehicle and in the case of emergency non-separation of the launch vehicle, after the carrier aircraft reaches at a preset time t
p
a flight path point with V
D
, H
D
, &phgr;
D
parameters and with a near-zero normal g-load, the launch vehicle is separated from the carrier aircraft and imparted with a speed relative to the carrier aircraft equal to the design speed at which the launch vehicle lags behind the carrier aircraft at a safe distance at the moment the launch vehicle sustainers are ignited, and before the launch of the launch vehicle to a scheduled trajectory point, the launch vehicle mated with the payload is turned to a position differing from the vertical by an angle of 10-30″ in the vertical plane in the launch direction.
The turn of the launch vehicle prior to its launch with the payload into a scheduled trajectory point is executed by sustainers after their ignition, or the turn of the launch vehicle prior to its launch with the payload into a scheduled trajectory point is executed by an additional jet engine prior to the sustainers' ignition. Separation of the launch vehicle from the carrier aircraft if effected, stabilizing the position of the carrier aircraft in an inertial coordinate system.
The object of the present invention is to enhance an aerospace system's lifting capacity when putting a PL into space, to ensure reliable separation of the LV from the CA, to provide safety for the plane and its crew members when firing the sustainers of the first stage of the LV, and to reduce the cost of developing an aerospace system and the cost of launching a pa
Borisov Andrei Vladimirovich
Ivanov Robert Konstantinovich
Karpov Anatoly Stepanovich
Kovalevsky Mikhail Markovich
Monakhov Jury Vladimirovich
Burns Doane Swecker & Mathis L.L.P.
Karpov Anatoly Stepanovich
Swiatek Robert P.
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