Structures Sybsystem Progress Report 11/03/95 - A&M

Structures Sybsystem Progress Report 11/03/95 - A&M

This week was spent researching deployment mechanisms (culminating in a visit to the Johnson Space Center on Friday to consult with engineers in the Structures and Mechanics Division), placing components into the spacecraft and determining approximate volumes available for vehicle subsystems. Figure 1 below shows how the vehicle fits into the fairing of the Delta-II (9.5 ft diameter option). The payload interface with the 3-stage Delta-II is achieved with the 3712C Payload Attachment Fitting. A payload attachment interface will be utilized that connects to a thrust ring directly below the propellant tanks on one end and has the 3712C PAF profile on the other end. The payload attachment interface transmits all thrust loads from the Delta-II to the lander structure and suspends the lander so that the lander's main engine does not interfere with the Star-48 upper stage. The nominal diameter of the 3712C PAF interface is 37 inches (94 cm) and tapers out conically to a diameter of 55.3 inches (140.5 cm) at a height of 38.6 in (98.0 cm). The 3 auxiliary thrusters (Marquardt R-4Ds) are arranged outside of this interface fitting on a 73.2 inch (186 cm) diameter at 120 degree intervals. The payload attachment interface ring will be jettisoned shortly after PAM-D separation. This figure also shows one possibility for landing gear deployment. It makes use of a hinging leg design that is deployed by the retraction of an actuator located in the lower strut of the leg. It also assumes that some type of load limiting device is being implemented, whether it is of the interference fit stroking washer variety or some other type.

Figures 2 and 3 below show the locations of some of the components in the vehicle. The primary payload (O2 plant and rover) is located in the top of the lander for the reasons described in the CDR. The payload is deployed via a ramp consisting primarily of two telescoping rails connected by either a series of rigid epoxy panels or by a fabric material such as Kevlar. The rover will aid in O2 plant deployment by pulling the plant down to the level platform connected to the end of the ramp. This will require minimal energy from the rover since it is going downhill and that the O2 plant is enclosed in a payload module equipped with small rollers on the base to minimize friction and ease in deployment. This payload module will have only two openings: the solar energy aperture used for collecting and focusing sunlight into the O2 plant reaction chamber, and the regalith hopper into which the soil sample is placed. The remainder of the O2 plant will be sealed in the payload module enclosure to provide adequate thermal protection. A small contingency sample collection arm will be mounted directly onto the payload module to allow for sample collection in the immediate area of the O2 plant in the event of rover failure. This arm will not be used at all unless the rover fails. Another approach was considered that consisted of a lander mounted O2 plant that featured a second sample collection arm that would be used to ferry the sample from the rover to the O2 plant and would double as a contingency sample collection device in the event of rover failure. However, this approach requires the sample collection device on the lander be on the order of 5 feet (152 cm) long and requires the use of two sample collection mechanisms to work in concert to get a single sample into the O2 plant. A lander mounted O2 plant also has visibility problems with getting enough sunlight into the solar energy collector since the 1.55 meter communications antenna essentially blocks all upward looking views. This would have to be routed around with a series of lenses and mirrors. It was felt that deploying the O2 plant in a thermally protected payload module to the lunar surface would alleviate these problems and provide a more fail safe operation since each sample collection arm operates independently of the other. Communication and power to the O2 plant will be provided by an umbilical tied into the lander's systems that unrolls from a spool as the payload is deployed. The payload module enclosing the O2 plant on the lunar surface can also easily facilitate a power coupling to mate with the rover in the event of rover power recharge.

Figures 4 and 5 below show the subsystem layout on the two primary component platforms (upper and lower as shown in Figure 3). The lower platform is located between the propulsion tanks and supports most of the spacecraft electronics. This centralized location helps protect the electronics from radiation upsets. The IMU is located as close to the CG as possible with the flight computers and other GNC electronics located adjacent to the IMU. The communications electronics (except for the primary antenna) are also located on the lower platform.

The upper platform (Figure 5) provides primary structural support for the payload (O2 plant and rover) during launch, transit, and landing as well as support for the power and thermal control subsystems as shown.

A study of this spacecraft configuration has led to the following updated values for the values available for each subsystem:
IMU and flight computer22.5" x 18.0" x 10.5" (57.2 cm x 45.7 cm x 26.7 cm)
Communications17.5" x 18.0" x 10.5" (44.5 cm x 45.7 cm x 26.7 cm)
(not including antenna)
Thermal Control35.0" x 13.5" x 13.5" (88.9 cm x 34.3 cm x 34.3 cm)
Power35.0" x 13.5" x 13.5" (88.9 cm x 34.3 cm x 34.3 cm)
O2 Plant11.0" x 26.0" x 22.0" (27.9 cm x 66.0 cm x 55.9 cm)
Rover24.8" x 18.9" x 11.0" (63.0 cm x 48.0 cm x 28.0 cm)
As always, let us know if there is a problem with these numbers.

A very rough estimate of the mass of the primary spacecraft structure is as follows (assumes graphite/epoxy strut construction with titanium lug joints and honeycomb panel platforms):

Total weight of graphite/epoxy struts20.3 kg
Total weight of titanium lug joints42.8 kg
Total weight of platform panels7.7 kg

Total structural weight70.8 kg

Note that this weight does not include the weight of the deployment and pointing mechanisms. This is also quite a bit higher than the value specified in the current mass budget (40 kg not including landing gear, robotic arm, or surface penetrators) but is a very coarse estimate with several conservative assumptions. A more detailed structural analysis will allow for a great deal of optimization and minimization of structural members and should reduce this weight considerably.

Moments of Inertia - Haven't determined this yet but should have a preliminary estimate this week.


Next Steps

- Develop designs for deployment/pointing mechanisms and refine solid model. The following mechanisms have been identified for development: 1) communications antenna 2-axis gimbal drive for pointing requirements, 2) payload ramp deployment mechanism, and 3) solar array deployment mechanism.


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