Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle Kevin E. Konno*, Daniel A. Catalano* and Thomas M. Krivanek* Abstract  As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft  Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is  based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust  vector control hard-over causing a severe vehicle pitch over. This worst-case scenario occurs just before  Upper Stage Main Engine Cut-Off when the vehicle is the lightest and the damping effect due to  propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed Simulink model to quickly assess the  Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach  involved the generation of an Automatic Dynamic Analysi s of Mechanical Systems (ADAMS) model to  assess secondary effects due to mass centers of gr avity that were slightly off the vehicle centerline. It  also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of  modeling results and accuracy are discu ssed. Most notably, incorporating a larger SA flange diameter  allowed for a natural separation of the Orion and it s engine nozzle even at relatively large pitch rates  minimizing the kickoff force. Advantages and disadvantages of the Simulink model vs. a full geometric  ADAMS model are discussed as well.  Introduction  A component of the Vision for Space Exploration, Orion will be capable of carrying crew and cargo to the  ISS, or rendezvous with a lunar landing module and an Earth departure stage in low-Earth orbit to carry  crews to the moon and, one day to Mars-bound vehicles assembled in low-Earth orbit. Orion borrows its shape from the capsules of the past, but it takes advantage of 21 st century technology in computers,  electronics, life-support, propulsion, and heat protection systems. Orion will be launched into low-Earth  orbit by the Ares I Crew Launch Vehicle. To maximize the crew’s safety, Orion and its abort system will be  placed at the top of the Ares I rocket. Other means of abort are available after the Launch Abort System  is jettisoned at ~75 km (250,000  ft). The Orion vehicle will be able to remain docked to ISS for up to six  months and have the ability to stay in lunar orbit untended for the duration of a lunar surface visit that  could be up to six months.  A separation mechanism design is being developed to assure clearance between Orion (Crew  Exploration Vehicle and Service Module) and the Spacecraft Adapter (SA), which stays fixed to the Ares  upper stage as the two vehicle elements separate from each other during both normal post-launch staging or in an abort event. Figure 1 depicts the Ares/Orion stack configuration prior to separation.   The preliminary design of the sep aration mechanism requires the balan cing of several competing design  parameters most notably sufficient kickoff forces to ensure separation, highly reliable components, limited  space to house these mechanisms, and a requirement to keep the mechanisms lightweight due to tight  mass budgets. The abort case will typically drive the size of the separation mechanism design for a crewed vehicle. Activation of the separation mechanism cannot occur until the thrust levels of the Ares I  Upper Stage (US) are significantly reduced. The potential hard-over gimbal abort case can induce a  severe pitch over rate (often referred to as “dump rate”) of up to 35  degrees/second on the stack if it were  to occur just before Upper Stage Main Engine Cut-Off when the vehicle is the lightest and the damping  effect due to propellant slosh has been reduced to a minimum. The transients of the controls and engine  thrust tail-off (~ 3 sec total from hard-over to low thrust) are the main reason a large dump rate can be induced. A Residual Engine Thrust also continues to induce a small force at 5 degrees off the vehicle  * NASA Glenn Research Center, Cleveland, OH   Proceedings of the 39th Aerospace Mechanisms Symposium, NASA Marshall Space Flight Center, May 7-9, 2008 345