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One of the most common launch vehicles for Earth-orbiting spacecraft is the Delta, which has launched over spacecraft and has low-orbit payload capabilities of kg to kg, depending on the rocket configuration.

For future launches, the Space Shuttle will be the most common, though not exclusive, vehicle. Launch sites are similarly limited. For discussions of Soviet space programs, see most issues of the British journal Spaceflight or the comprehensive Senate report for the 92nd Congress Once powered flight has ended' and the spacecraft has separated from most of the launch vehicle, the acquisition phase of maneuvers and testing begins. The final launch stage may be left attached to the spacecraft for later maneuvers using the final stage control hardware and fuel.

An Agena upper stage, a rocket which may be restarted in space, is frequently used for such maneuvers. The acquisition phase can last from a few minutes to several months and may be defined differently depending on the particular aspect of the mission that is involved. For example, for someone concerned primarily with the operation of communications hardware, testing and maneuvering may last only a brief period, with "normal" operations beginning well before experiments or operational equipment have been thoroughly tested.

Normally, the major portion of the attitude analysis for any mission is concerned with various aspects of the acquisition phase, as described in the example below. Finally, once the proper orbit and attitude have been obtained and the hardware has been tested, the mission operations phase, in which the spacecraft carries out its basic purpose, is initiated.

At this stage, attitude determination and control becomes, or should become, a routine process. On complex missions, such as lunar or planetary explorations, the acquisition phase may be repeated at various intervals as new hardware or new conditions are introduced.

We will illustrate the above phases and the role of the attitude determination and control process by describing the flight of the Communications Technology Satellite, CTS, launched aboard a Delta launch vehicle Fig. CTS was a joint project of the United States and Canada in which the Canadian Department of Communications built and operated the spacecraft and the United States National Aeronautics and Space Administration, NASA, provided the launch vehicle, launch facilities, and operational support through the acquisition phase of the mission.

The spacecraft was placed in synchronous orbit, i. This was done to avoid the van Allen radia ti on belts. Conditions for the launch of the Synchronous Meteorological Satellite shown here were essentially identical with those for the CTS launch which occurred after dark.

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This location permitted 'television transmissions to remote regions of both Canada and Alaska. The mass of the CTS spacecraft was kg at lift-off, of which approximately kg was in the weight of a rocket motor, called the apogee boost motor, required for an orbit maneuver during the acquisition phase.

As shown in Fig. The main operating power is supplied by two extendable solar arrays, each 6. CTS has a total of 11 attitude sensors': In addition to the large apogee boost motor used during the acquisition phase, the spacecraft includes 18 small rocket motors 2 "high thrust" and 16 "low thrust" for orbit and attitude maneuvers. See Chapter 3 for a general discussion of orbits and orbit terminology. Before injection, which marked the end of the launch phase, the spacecraft was controlled by the launch control team at the launch site.

Orbit and Earth drawn to same scale; see text for explanation. The purposes of the transfer orbit were to move the spacecraft to synchronous altitude so that the large apogee boost motor could be fired to change the orbit to an approximately circular, synchronous one, as shown in Fig. The smaller motors could then be used for various orbit and attitude refinements.

Because the apogee motor was fixed in the spacecraft, it was necessary to reorient the entire spacecraft such that the motor firing would provide the proper orbit change.

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Thus, the p ri ncipal activity during the transfer orbit was to determine the attitude, test and calibrate if needed the attitude sensors, reorient the spacecraft to the proper apogee motor firing attitude, and make fine adjustments and measurements as needed to ensure that the proper attitude had been obtained.

While in the transfer orbit, the spacecraft was spin stabilized at approximately 60 rpm. Two Sun sensors and two Earth horizon sensors were used for attitude determination with an accuracy requirement of 1 deg for apogee motor firing. The two high-thrust rocket motors were used to reorient the spacecraft spin axis by about deg from its initial o ri entation as it left the third stage of the launch vehicle to the apogee motor firing attitude.

A maneuver of mo re than deg was required to avoid lowering perigee the minimum altitude point in the transfer orbit due to the translational thrust of the maneuver jets. The apogee motor firing changed the orbit to nearly circular and changed the period to approximately 23 hours 15 minutes so that the spacecraft would d ri ft slowly westward relative to the Earth's surface. A series of orbit adjustments made the period nearly identical with the Earth's rotation period when the spacecraft was over the desired longitude.

The station acquisition sequence required a total of five progressively smaller orbit maneuvers carried out over a period of 9 days after apogee motor firing. After station acquisition, control of the spacecraft was transferred to the Canadian Research Council, which conducted a major attitude maneuver sequence to transform the spacecraft from a spin-stabilized mode to a three-axis-stabilized, Earth-pointing mode [Bassett, ]. Th is was the most complex attitude maneuver conducted during the mission and consisted of a 2-day sequence of operations divided into 39 specific events.

During this phase, attitude determination input was changed from the four sensors previously mentioned to five Sun sensors and two Earth sensors designed for use during the acquisition phase.

The set of 16 low-thrust jets was used to despin the spacecraft and to control it in the nonspinning mode. Additional control was supplied by a spinning flywheel, which was used to reorient the spacecraft about the wheel axis by changing the relative angular momentum of the wheel and the spacecraft body. Completion of the attitude maneuvers ended the acquisition phase of the mission. After further hardware checks, normal mission operations were initiated. During the planned 2-year life of the spacecraft, the attitude control system will be used to maintain the attitude within 0.

The major factor in mission lifetime is the consumption of fuel for attitude stabilization, although orbit dri ft, possible mechanical failure, and power loss due to radiation damage to the solar cells may also affect the useful life of CTS. During the decade of the s, three major spaceflight changes are anticipated which will affect the representative mission profile just described: The effect of these and unforeseen developments on attitude determination and control hardware and procedures cannot be predicted with precision.

It is nevertheless important to consider the probable or possible direction of future developments. The potential cargo mass for various shuttle orbits is shown in Fig.

Spacecraft Attitude Determination and Control (Astrophysics and Space Science Library)

As transportation costs decrease and the number of active payloads increases, new methods will be needed to reduce the cost of attitude determination and control. At present, the most likely procedures to achieve this are 1 increased autonomy with up to 3 days of automatic control without ground support; 2 decreased hardware redundancy, since recoverable payloads will shift cost effectiveness; and 3 standardization of attitude hardware and, possibly, supporting software: Broken Circle by BmblBee Tara has been abducted and the police rush to find her.

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