The Delta II-7925 was selected as the launch vehicle for this mission. This confirms the preliminary choice from last semester. It was chosen because it provides a great deal of energy for lift-off, including nine bosters, six of which ignite at lift off and stage at 57 seconds. The first and second stages use inertial guidance and provide the control moments gimballed engines. The third stage is derived from the components and the concepts used on the delta third stage and the USAF SGS-II upper stage. The Star- 48B, solid rocket is supported at the base of the motor on a spin table that mates to the top of the second stage guidance section. The payload attach fitting is the structure that provides the transition from the top of the solid rocket motor to the space craft interface. The payload fairing shields the payload from buffeting and aerodynamic heating while in the atmospheric phase of flight. Payload Constraints During the burn periods, a pressure differential occurs because the ambient atmospheric pressure continuously drops with altitude while the fairing contains the higher pressure air. Air trapped in the compartment and crevices within the fairing and the spacecraft itself is at a higher pressure until it is vented to the outside. The venting rate depends on the pressure differential between the internal payload compartments and the fairing. It is also dependent upon the pressure differential between the volume enclosed within the fairing and the external environment. The payload components must be vented to prevent damage during ascent. For those reasons the chamber pressure contains 39.7 bars, or 575 psia, and the diameter of the nozzle exit is 2.0 m. Guidance The digital inertial guidance system is mounted inside a stage-2 cylinder on the forward end. That controlls the vehicle during the flight and commanding spin-up of stages 1 and 2, and separation of stage 3. It also is used to trigger its fuze-based sequencing system. The computer also issues preprogrammed sequence commands and provides attitude control. The payload fairing is the aluminum structure, which incorporates acoustic absorption blankets on its interior and accommodates the space craft envelope. The volume of the payload section is 2.5m in diameter by 2.0m cylindrical height and 2.9m conical height. The acceleration load is at a maximum of 5.77g at stage 1 burn associated with the heaviest GTO (geosynchronous transfer orbit) payload; 4.24g at stage 3 burn out; and 1.31g at launch. The acoustic load is at a maximum of about 130 dB at launch and transonic. The temperature is at its maximum of 33 C on the fairing inner wall after 30 seconds. Payload integration, nominal mission schedule begins, is 48 months. Environment, maximum load factor +6.0g axial, +- 2.0g lateral, minimum lateral and logitudinal payload frequency is 15 hz to 35 hz. The maximum dynamic pressure on the fairing is 1230 lb/ft2 (58898 N/m2). Payload delivery, standard orbit and accuracy (3s) LEO (low earth orbit) +-18 km, +-.5 deg. inclination; GFO (geosynchronous transfer orbit) +-6 km perigee, +-740 km apogee +-0.2 deg inclination, Altitude accuracy (3 sigma) +-1 to 2.5 deg. Nominal payload separation rate 0.3 m/s. The available deployment rotation rate is 20 rpm. To separate the launch vehicle from the spacecraft pyrotechnique devices are typically used. These devices are light, highly reliable, and easily integrated into mating techniques that provide a high degree of stiffness.

The design selected for the Main Propulsion System (MPS) is a system with one main engine and three auxiliary thrusters. The main engine will be responsible for the three major burns plus correction maneuvers and a de-orbit burn. The three thrusters will have the capability to hover the lander and descend the lander to the surface of the moon. The mission is divided into ten phases. These phases and their required DV's are listed below. The tasks for the main engine start from the earth departure burn to the first descent phase. At this point the main engine is jettisoned and the three auxiliary thrusters take over. The thrusters will hover the vehicle and land it on the moon. Major Mission Phases DV's (m/sec) Earth Departure Burn 3108 TCM #1 20 CM #2 20 TCM #3 20 TCM #4 20 Lunar Orbit Insertion 825 TCM #5 20 De-Orbit Burn 20 First Phase of Descent 1665 Last Phase of Descent 20 Engine Selection: The main engine selected is the Rocketdyne XLR-132 and the three thrusters are MR-107. Specifications are listed below: ROCKETDYNE XLR-132 MR-107 Dry mass: 54 kg .885 kg Length: 120 cm 21.8 cm Max Dia: 60 cm 6.6 cm Thrust: 16.68 kN .178 kN Isp: 340 sec 220 sec Burn time: 4000 sec., 10 starts 150 sec

The Conceptual Design Report Specified two different proposed designs. The first was a design consisting of an upper and lower spherical manifold mounted with numerous thrusters, directed normal from the surface. This was believed to be exotic. The second design specified a coupled pair of thrusters for each perpindicular direction on each axis and for each stage. We have recently been told that it is not necessary to couple the thrusters, and that the translation that would occur would not be significant. Therefore, the conceptual design has been modified to having only one thruster for each perpindicular direction on each axis. This require total six thrusters, two for each axis. To calculate the thrust needed the lander was assumed to have a uniformly distributed mass. Its mass, immediately after detachment, was assumed to be 800 kg. For the stages proceding the main engine being jettisoned a mass of 350 kg. was assumed. These are estimates we assumed based on our current mass calculations. They will change as the numbers become more solid. The thruster locations were assumed to be as far away from the center along each axis as possible. That was equal to the radius of the lander. The radii about the x- and z-axes were assumed to be 1.0 m. The radius along the y-axis was 1.45 m. It was assumed to be 0.85 m after the main engine was jettisoned. This was due to the fact that the length of the XLR-132 was 120 cm. These masses and radii were used in the equation F = m * a * r The necessary thrusts calculated about the x- and z-axes were equal. The calculated thrusts prior to the jettison of the main engine were 607 N, and after the jettison, 132 N. About the y-axis the thrust before the jettison was 418 N, and after the jettison it was 188 N. These thrusts are to be used to size the thrusters to be used in the Reaction Control System. The only reference we had from which to chose thrusters was the Interavia Space Catalogue. It contained no thrusters with a specified range as we needed. Therefore, we will have to consult further references.

As for the lander, we need to determine the final orbit parameters. The main propulsion burn times, fuel and oxidizer requirements, and tank size must be calculated. Te reaction control fuel type, burn times, and mounting structures must be determined, and the rotational inertias must be acquired.Questions: What other references can we use to look up thrusters?

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