CAD Diagram of the Spacecraft Structure
This page describes the SNOE spacecraft structural design and component
layout. To meet the low cost objective and the science requirements a spin
stabilized spacecraft was chosen. The structure consists of two hexagonal
solar arrays that attach to a central mounting plate. The mounting plate
supports the spacecraft electronics and equipment, as well as the scientific
instruments. These units mount to both sides of the mounting plate. Two
patch antennas protrude slightly above the ends each solar array. Attached
in a band to the periphery of the central support plate, between the solar
arrays, are six thermal radiator plates. These plates have apertures, through
which the instruments, and two horizon crossing indicators, observe. Thermal
blankets cover the open hexagonal ends of the spacecraft.
A key consideration in the construction of the spacecraft is to configure
the payload to produce spin stability . Trade studies were performed on
the spacecraft design to arrive at the present design. Orientation in orbit
was balanced against power requirements (and solar cell collector area)
to arrive at this final layout. In each design iteration the payload height
was shortened, and the mass was distributed to the outermost locations
on the mounting plate. The design shown above is spin stable, with spin
to transverse axes ratios of 1.2, significantly larger than the 1.05 necessary
for spin stability.
A second consideration is the mass of the spacecraft structure. The
tradeoff on this issue balances cost, mass, and ease of fabrication. Metals
used in the assembly are principally aluminum, with some aluminum honeycomb
(solar panels), and titanium (thermal flexures). The final weight of the
fully integrated spacecraft and instruments is 254 pounds.
The structural solution to the spacecraft design involves four primary
structural components: 1) the spacecraft equipment and instrument mounting
plate; 2) the adapter structure that mates the spacecraft to the launch
vehicle. This structure has provision for a separation assembly (marmon
clamp and actuator) to release the spacecraft on orbit after positioning
and spin-up; 3) two solar array assemblies and their support structure;
and 4) two antenna masts.
The central mounting plate provides support for the spacecraft electronics,
instruments and cables, and the launch adapter structure. It mounts both
of the solar array structures and the antenna mast assemblies. It provides
also a thermal path for heat generated in the interior of the spacecraft
to reach the thermal radiators. The central mounting plate is fabricated
from two weight-relieved alumnium plates bolted together in a clamshell
The launch vehicle adapter structure mates the spacecraft to the launch
vehicle. The assembly consists of twelve struts (1 inch diameter and .083
inch wall thickness) which connect the central plate to a 23.25 inch diameter
marmon clamp on the launch vehicle end. The struts are attached to the
marmon clamp and central plate in pairs through connection blocks machined
out of aluminum. The separation fixture details have been developed in
accordance to the details outlined in the Payload Users Interface Control
Document (ICD). During integration the adapter structure will be attached
to the launch vehicle's marmon clamp using a clamp band.
The two hexagonal solar array assemblies consist of 6, 0.5" thick
honeycomb panels that are approximately 18.75" wide by 13.5"
tall. Mounted to each panel are two strings of solar cells; the cells are
bonded to the surface of the honeycomb material using traditional mounting
techniques. Two edges of each honeycomb panel attach to axial supports
which serve as columns providing axial (spin axis) rigidity. Each set of
six assembled solar arrays is a monocoque, using the panels as shear ties.
The assembled structure joins to the spacecraft by bolting the axial supports
to thermal flexures which in turn are bolted to the central mounting plate.
The thermal flexures consist of a thin (0.1 inch) "blade" of
titanium 0.5in in length. This approach allows the spacecraft equipment
to be managed thermally without concern for the varying solar panel temperatures.
The construction of the top and bottom solar array structures is essentially
identical. Each end of the spacecraft requires a closeout of MLI blanket
and Beta cloth for thermal reasons. On the bottom end the spacecraft adapter
structure will provide the necessary support for this material; on the
top end of the structure a "spider" is shown to provide the necessary
tie downs and fastening locations.
There are two patch antennas located on the spin axis at each end of
the spacecraft. The antennas consist of a support tube and a thin plate.
This assembly mounts to each side of the central mounting plate. The column
provides structural support for the antenna, and precisely positions it
so that the ground plane of the antenna sits above the solar panel.
Attached to the central mounting plate, around the periphery of the
hex, are six 0.1" aluminum radiator panels. Three of these panels
have apertures for the science instruments and sensors located on the spacecraft.
The radiators are coated with a durable ceramic surface to maintain radiative
properties so that operating temperatures are controlled.
A finite element model of the spacecraft structure was constructed using
Cosmos/M. The model consists of 125 tri-shell, quad-shell and 3 layer composite
elements. Modal analysis for the first ten modes were analyzed, and simulations
were run to determine mode response and displacements. The structural model
includes the adapter structure, the mounting plate and solar arrays. Masses
representing most of the spacecraft electrical and sensor components, as
well as the spacecraft instruments, were also included. The first mode
frequency is 40 Hz, and corresponds to a cantilever mode shape off the
launch adapter marmon clamp assembly.