Deploy ment Mechanism for the Space Technology  5 Micro Satellite 

Peter Rossoni *, Caner Cooperrider * and Gerard Durback*    

Abstract 

Space Technology 5 (ST5) is a technology mission that  will send three spin-stabilized, 25-kg satellites  into a highly elliptical Earth orbit. Each of these satellites must be deployed separately from the same  launch vehicle with a spin rate of 3.4 rad/s (32. 4 rpm). Because of the satellite’s small size and the  requirement to achieve its mission spin rate on deploy, typical spin table, pyro technic deployment devices  or spin up thrusters could not be used. Instead, this  new mechanism design employs a “Frisbee” spin up  strategy with a shape memory alloy actuated Pinpu ller to deploy each satellite. The mechanism has  undergone several design and test iterations and has been successfully qualified for flight. 

Introduction 

The ST5 mission has several important purposes. The first is to integrate new scaled down technologies  into a miniature spacecraft constellation operating in an environment typical of future magnetospheric  mapping missions. The second is to serve as a pathf inder for development, integration, and deployment  of larger constellations  of miniaturiz ed spacecraft. The final purpos e is to tak e highly ac curate s cience  measurements. The ST5 three-small-satellite configurat ion was originally slated to launch as a secondary  payload on an Atlas V, Delta IV, or Ariane 5 expendable launch vehicle (LV). Currently, ST5 is a likely  candidate for launch on a Pegasus in late 2005. As a re sult of this LV uncerta inty, the ST5 spacecraft  bus, deployer, and mechanisms faced the added challenge of being designed to survive the worst-case  launch levels and envelopes for all of these possible LVs.  

Initial Deploy er Design 

The ST5 deployer was initially designed to sati sfy a number of requirements. These are:  • Deployer must deploy the spacecraft with a spin ra te of 3.4 rad/s (32.4 rpm) +/- 10% with tip-off  Nutation < 10˚. The mission spin rate of 2.6 rad/s (25 rpm) is achieved by de-spin upon boom deploy.  • Deployer and spacecraft must survive worst-case LV  loads applied uniaxially in all three directions.  These include 17.5 G acceleration, 12.0 Grms r andom vibration, and 3.75 G sine sweep testing.  • Deployer and spacecraft must fit inside the worst case LV envelope of a 60-cm cube.  • Deployer and spacecraft must have a natural frequency of at least 35 Hz.  • Deployer and spacecraft must minimize  the use of magnetic materials.  • Deployer and spacecraft must survive ST5’s ascent and early orbit temperatures of –55 oC to 60 oC.  • Deployer and spacecraft must function in  its operating temperature range of -20 oC to 50 oC. 

The ST5 Deployer designed to satisfy these requirem ents is shown in Figure 1. The deployer base is  7.62-cm (3 in) thick aluminum honeycomb sandwiched by 1.52-mm (.06 in) thick aluminum face sheets.  Co-cured inserts accept the twelve NAS6705U54 bolts that interface to the LV and also provide rigid  attach points for the three machined aluminum st anchions in the corners and for the four machined  aluminum braces that laterally support these stanchi ons. The spacecraft bus is supported in the deployer  at the three mechanisms at the tops  of these stanchions. The spacecra ft interface is semi-kinematic.  

The spacecraft is fixed in its Y and Z directions by the YZ mechanism shown in Figure 2. The 455stainless-steel pin of the shape-me mory-alloy-actuated Pinpuller on the depl oyer fits into the aluminum  clevis of the spacecraft’s YZ hardpoint. This YZ hardpoint clevis accommodates a TufLite® bushing with a  clearance around this pin of 0.051 mm (0.002 in) to 0.102 mm (0.004 in).                                                     * NASA Goddard Space Flight Center, Greenbelt, MD 

Proceedings of the 37th Aerospace Mechanisms Symposium, Johnson Space Center, May 19-21, 2004 

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