Propulsion Group Progress report 10/27/95

Propulsion Group 10/27/95

This is the progress report for the Lamar Propulsion Group for the week of October 27. For this week the following work was done.

Launch Vehicle Selection

(Diep Nguyen)

        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.

        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

Selection of Preliminary Design and Specific Engines

(Britney Sandell)

        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         

Reaction Control System Thrusters Selection

(Gary Rhodes)

        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
        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.

For the next week:

        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?


Wednesday, 31-Dec-1969 18:00:00 CST
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