ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
pressure loss. Based on the analytical results, one strut failure cannot result in the loss of an Ares I  mission due to nozzle contact. Consequently, the MAG Strut system is inherently two-fault tolerant.   Plume Heating on Upper Stage  At the Ares I System Definition Review, the vehicle was configured with BDMs mounted near the aft end  of the interstage in four pods containing two motors each. The USMs were mounted on the upper stage aft skirt, also in four pods of two at the same angul ar positions around the cylinder. One problem with this  configuration is the interaction of the USM and BDM plumes. Even though the nozzle exit planes were separated by over 4.5 meters axially, extreme heating was predicted in the upper stage engine  compartment during separation because the BDM plumes deflect the USM plumes into the interior of the interstage. Also, debris generated by the separation pyrotechnics will likely be propelled into the engine compartment by the interacting plumes. The use of gas struts eliminates these debris and heating  concerns. Relocating the BDMs to the first stage aft skirt would resolve this issue.   Payload-to-Orbit Benefits   Gas strut separation produces a significant increase in payload-to-orbit capability. This gain is a result of  reduced aerodynamic drag, momentum transfer between the stages, and ascent trajectory optimization.  The interstage-mounted BDM pods are the largest protrusions from the nominal outer moldline (OML) of the vehicle. As such they account for a total of a 110 to 120 kilogram payload pe nalty due to aerodynamic  drag. The proximity of the BDMs to transition from the conical to cylindrical is a major factor in the high drag. Locating the struts inside the interstage eliminates all aerodynamic drag effects.    For the baseline trajectory, the amou nt of residual first stage thrust at separation is limited by the  capability of the BDMs. For an 8 BDM configuratio n with one motor out, separation must wait until first  stage thrust drops to 178 kN. Because the struts have a better weight to performance ratio than BDMs, the trajectory can be optimized to improve performance. Figure 5 indicates the amount of payload that can be gained relative to the baseline flight profile. The steeper section of the curve indicates a significant payload improvement, but the strut system mass (including additional upper stage structural mass) begins to offset the benefit as residual thrust increases. Sepa ration at 356 kN  of residual first stage thrust  is thought to be optimum for Ares I. This results in approximately 90-kg additional payload due to improved trajectory performance. 

Figure 5 - Payload Delta from Baseline vs. Separation Thrust  334
During separation with gas struts, the first stage thrust continues to act on the upper stage until the end of  the stroke. Initial calculations show that this momentum transfer adds payload performance at a rate of 8.93 kilograms for every meter per second of /g507V. Preliminary strut designs result in an increase in upper stage /g507V of 3 to 3.7 meters per second. This amounts to 27 to 33 kilograms of additional payload. Figure  6 shows the relative velocity gained by the upper stage  for a separation with 356 kN of residual thrust. 

Figure 6 - Momentum Transfer Effects for a Separation at 356-kN SRB Thrust   The mass of the struts and upper stage fittings for a 356 kN thrust separation are about half that of a  BDM system that separates at 178 kN of residual thrust; however, because more of the mass remains with the upper stage, no additional payload advantage from the change in system mass is realized.  Table 1 - Approximate Payload Benefit  Reduced Drag 110 kg  Earlier Separation 90 kg  Momentum Transfer 27 kg  Mass Delta Benefit 0 kg  Total Payload Benefit 227 kg 

Cost Considerations   The projected unit cost for each BDM is approximately $200,000. There are many reasons for this high cost. One of the most risky processes of solid rocket motor manufacturing is the casting and curing of the solid rocket propellant. The process is very hazardous and requires extensive risk mitigation to prevent inadvertent propellant ignition. The risk mitigation techniques are well known, and accidents are now rare, but the process is expensive. Additionally, post-casting inspection sometimes reveals defects in the cast  propellant. If a defect is found, most often the motor is discarded.   Per unit cost for gas struts should be significantly less than BDMs, since there is no hazardous material to procure and handle. Also each flight unit can be acceptance tested, so manufacturing will not require the  strict process control necess ary for solid motors. If a defect is discovered during the acceptance testing,  335
the total exposed hole-area for two different hole-sizes in shown at the bottom of the chart.) A large range  of force profiles is possible with different hole-patterns. Holes larger than the “O” ring seal diameter would likely catch the seal, causing damage during stroking. A hole diameter of 3.96 mm would be the largest recommended hole size for a seal with a 4.83-mm diameter cross-section.

Figure 8 – Average Thrust and Oscillatory Thrust Test Data for 5 Segment SRM 

Figure 9 - Strut Rendering (Collapsed) 

Figure 10 - Strut Rendering (Extended) 

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Figure 11 – Exposed Hole Area for Two Candidate Hole-Patterns  If the low pressure chamber is allowed to be at ambient pressure by providing a very small hole to the  exterior of the strut, the strut can operate with only one pr e-pressurized volume. This variation would  make it possible to charge only on e chamber prior to launch, elimin ating some potential failure modes.  A  strut with a 9.208-cm diameter metering rod and with no pressure in the small chamber would provide slightly more initial separation force th an the strut shown in Figure 7. This strut variant opens up the  possibility of designing a hermetically sealed strut or other point design.   Since the desired thrust profile for the struts is based on requirements de rived from a fluids analysis of  the hydrogen tank pressure, having a strut capable of accommodating a range of force profiles is preferable. For a -147 degree C  initial ullage gas charge temperature, an acceleration rate of change of  2.5g per second is acceptable. A higher axial rate of change may be acceptable with the currently proposed -220 to -250 degree C pre-charge gas. Table 2 shows the predicted effect of lowering pre-charge gas temperature on the make-up gas required to recover from an ullage collapse. A change out of metering rods could adapt a set of struts to revis ed ullage requirements. Sloshing risk increases as the  axial acceleration of the rocket diminishes. Surface tension and vibration force the fluid in  the tank up the  tank walls as shown in Figure 12. Stage separation with 356 kN of residual thrust assures that the average axial acceleration never drops below .12g. This is enough acceleration to force the ullage gas to remain in a hemispherical shape bubble. The MAG Strut system further mitigates the risk of ullage collapse by limiting the axial acceleration rate of change.    Table 2 – Hydrogen Tank Recovery Gas Requirements   Initial tanked He assumptions: T=-250 C; P=22,00 kpa                       Supply assumption: Isentropic Blowdown P=6,895 kpa   H2 pre-press  temp Mass for  ullage recovery  Initial storage  density  Final storage  density Delta density Storage volume  requiredBottle  massTotal  loaded  He mass Total  Mass 19 C  -181 C -220 C -250 C 226.9 kg  115.7 kg  0.00.0192.38 kg/m 3 192.38 kg/m3 192.38 kg/m3 192.38 kg/m3144.81 kg/m3 144.81 kg/m3 144.81 kg/m3 144.81 kg/m347.58 kg/m3 47.58 kg/m3 47.58 kg/m3 47.58 kg/m3.486 m2 .248 m2 0.00 0.001,390.0 kg  708.6 kg  0.00.0917.2 kg  467.5 kg  0.00.03,642.2 kg  1,176.5 kg  0.00.0

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Figure 12 - Ullage Gas Collapse Mitigation Proposals   Real fluid analytical tools show that the smaller holes produce a force-profi le that does not exceed  8,896.4 kN  per second level as shown in Figure 13. The force-profile has some irregularities that can be  eliminated through further refinement of the hole-pattern. T he force spike at .4 seconds indicates that a  few more holes are needed in the last 7.62 cm of sto ke for the 3.18-mm  diameter holes. If the first row of  holes were exposed after 1.27 cm of stroke rather than 2.54 cm of stroke, more energy could be recovered from the expanding gas. If a few less holes were exposed in the middle part of the metering  rod, the rate of change peak could be lowered. For Ares I, the 3.18-mm  diameter holes shown in this plot  meet a 2.5 g/sec jerk requirement if the decay of the thrust of the SRB is considered.  

Figure 13 – Force Rate of Change Plot   339
testing will establish the amount of side loading required to cause the strut to bind such that the load relief  action from the conical interfaces will not be adequate to relieve it.                 Figure 17 – Strut Fitting Aft Skirt Mounted Spike Fitting  

Resolving Reliability Related Concerns    The safety of the struts must be demonstrated by test. The struts are designed to leak before burst; however, only testing can demonstrate this. If the leak before burst design is proven prior to qualification, the potential for a costly redesign and schedule slip is avoided. After completion of testing, one or more of the test struts would be subjected to extreme pressure until leakage or burst occurred. This burst test would be done with an oil or water charge to avoid the explosive hazards associated with gas.    All elastomeric seals leak a minute amount of gas because of permeation of the seal material. The expected performance of each seal must be bounded in order to establish launch commit requirements and pad operations. Nominal leak rates of the seals could be established without assembly into the struts by using a test fixture as shown in Figure 18. Different elastomer compounds could be evaluated for gas permeability at the pressures used in the strut. With this data the struts could be pressurized taking into account the number of days before launch. The low pressure chamber would gain a very small amount of pressure due to seal permeation during pad operations but not enough to exceed its required operating range.   Pressurizing the large volume chamber while leaving the low volume chamber at ambient pressure as discussed in the performance section of this paper would also be an option to eliminate uncertainties about rate of leakage into the low pressure chamber from the high pressure chamber. Figure 18 shows potential test configurations for two different seals. Testing 50 seals of each type would provide a large enough sample size to characterize the nature of the seals under ambient conditions. Temperature extremes could also be evaluated by placing the small seal test fixture in a thermal chamber.    Analysis Needed Prior to System Testing  An analysis of the integrated system would be required to establish the overall capability of the MAG Strut  system to achieve separation under all potential operational scenarios. Initial analysis shows startling  results with large positive clea rance margins for th e nozzle during separation. Revisiting this analysis is  required prior to system testing to assure that an undiscovered disturbance force acting in the system will  not cause the results to degrade. 

To recover the first stage, the interstage with the extended struts must be separated from the first stage. However, no analysis has been done to establish the clearance between the first stage and the  interstage. The struts extend about 2.44 meters from the interstage. Consequently, their presence will  make it more difficult to gain adequate clearance between the first stage and the interstage after separation of the interstage from the first stage.    342
Figure 18 - Proposed Seal Test Configurations 

Stress analysis of the second stage aft skirt interface with the spike fitting would provide a better  understanding of the threat of buckling with a failed strut. If the high-pressure seal fails on a strut, the good strut will apply 68 kN of load to the structure  while the failed strut will apply 236 kN of load. The  safety factor is 1 for analyzing a failure case.  However, the safety factor is 1.65 for buckling without a  failure. Showing sufficient margin under all conditions is required prior to approving a final design  configuration.   

A stress analysis using finite element models of the struts themselves is required to assure adequate margin exists for all components. This analysis would allow for weight optimization of the strut prior to  finalizing the design. 

Integrated System Testing 

Testing the integrated system has the decisive advantage of establishing the validity of the analytical  models used to evaluate separation dynamics. A close match between the development testing and the  analytical models will make it possible to qualify the separation dynamics by analysis, avoiding an expensive flight test dedicated to qualifying the separation system. Actually simulating the flight conditions  is not practical considering the cost and complexity of such a test set up. A test setup that is capable of  simulating any flight condition in one plane could be used to demonstrate the system incrementally. Figure 19 shows a proposed test setup that would be capable of simulating all of the most relevant  conditions in the horizontal plane.  

Figure 19- System demonstration test set up  

Different asymmetric strut cases could be combined with various simulated thrust conditions. The  simulations could be accomplished by placing many support points at the center of gravity of each of the  mass simulators. The brake rod would have a ball joint attachment at the center-of-gravity and the brake  body would be free to rotate on a pivot arrangement. When the separation joint is activated, the brakes  would simulate the effects of the SRB thrust and the relevant component of gravity acting on the vehicle. This set up would simulate the mass and the mass-moment-of-inertia of each of the stages. Thrust vector  side loads would be simulated by springs acting between the rod coming from the brake and the end of  the first stage. The brakes would also arrest the motion of the two bodies after separation was  343
centerline resulting in an applied moment to the Upper Stage after separation. This defines the worstcase environment that the separation mechanisms must overcome. Recontact during either an abort  separation, or a nominal separation, can be catastrophic resulting in a Loss of Mission or a Loss of Crew  event. A recent example of this type of detrimental recontact  was observed during  the March 2007  SpaceX Corporation Falcon I launch first stage separation event in which the first stage adapter inner wall  contacted the second stage engine nozzle and induced a propellant slosh in the second stage tanks,  prematurely shutting down the second stage engine before reaching the proper orbit [1]. The Orion separation system must be adequately sized to reliably separate the crew and vehicle safely for all design  cases. Multiple types of mechanisms were evaluated including spring actuators, constant pressure  pneumatic actuators, and pyrotechnic-actuated gas thrusters.  Crew Exploration  Vehicle (CEV) Solar Arrays (2) Spacecraft Adapter (SA) Ares 1 Launch Vehicle J-2X engine & remainder of vehicle shown on leftService Module (SM) Service Module Engine NozzleAres I Upper Stage (CLV) J-2X EngineAux Engines (8 total) Crew Exploration Vehicle (CEV) Solar Arrays (2) Spacecraft Adapter (SA) Ares 1 Launch Vehicle J-2X engine & remainder of vehicle shown on leftService Module (SM) Service Module Engine NozzleAres I Upper Stage (CLV) J-2X EngineAux Engines (8 total) Crew Exploration Vehicle (CEV) Solar Arrays (2) Spacecraft Adapter (SA) Ares 1 Launch Vehicle J-2X engine & remainder of vehicle shown on leftService Module (SM) Service Module Engine NozzleAres I Upper Stage (CLV) J-2X EngineAux Engines (8 total) Figure 1. Ares I Upper Stage/Orion Spacecraft Conf iguration (Lockheed Martin Concept)  The Point of Departure separation system is shown in Figure 2. Unlike Apollo’s Service Module, which  was bolted to the top of a four piece faring and severed from it via a circumferential linear shape charge,  this system incorporates compression kickoff springs and pyrotechnic separation bolts to join the Orion Vehicle to the SA. Separation is triggered by firing the pyrotechnic retention bolts, which allows the  compression springs to push the SM away from the Upper Stage. The spring force must be sufficient to  accelerate the separated bodies away from each other while maintaining a minimum clearance throughout separation.  346
Figure 2. Separation Spring Concept  Separation System Hardware The force needed for separation can be generated from various competed technologies, including  mechanical springs, pneumatic actuators or gas thrusters. The other components of interest in the  staging mechanism are the pyrotechnic fasteners through which the launch loads are transmitted.   Mechanical Springs: For this study open coiled, helical, compression springs were the preferred form of  kickoff devices if the required energy level was low enough to warrant their use. Spring kickoff devices  were incorporated in most of the models because the separation environment did not require large kickoff  forces. Obviously, mechanical springs are used in countless terrestrial applications as well as space.  They are highly reliable and when designed correctly can handle millions of cycles. Several papers detail the use of compression springs in spacecraft staging mechanisms [2], [3], [4], [5], [6], [7], [8]. For  spacecraft mechanisms mechani cal kickoff springs must be designed  with the resistive force (F r) and the  force required for acceleration (F a) of the bodies in mind. Where redundant springs are used instead of a  backup mechanism, they should be designed to provide adequate force for a one-spring-out case [5].   Additionally, the spring system should have a 100% positive Margin of Safety on drive force over resistive  force, as measured at acceptance or qualification testing. It is prudent to carry additional margin prior to  testing. Also, spring systems are required, when practical, to have a dynamic force margin of safety over  the required force F a of 25%, as tested. Additionally, when sizing mechanical springs, spring material  stress relaxation and resid ual stresses must be factored in, for which some test data exists [9]. These  effects decrease the driving force a spring is capable of after prolonged storage, and can vary by greater  than an order of magnitude depending on the material. 302 SS, a common aerospace spring material, can go through 3-5% stress/preload relaxation in 1000 hours of storage time.   The dynamic modeling of springs in this system always considered six compression springs located  equidistant around the circumference of the SA interface flange. If incorporated in the final flight design,  equivalent redundant pairs of springs will be used to improve reliability. It is important when designing  mechanical redundancy to do so wisely as it has been shown that some redundancy can actually decrease overall system reliability, even in spring actuator designs [10], [11]. 347
Pneumatic Actuators: Pneumatic actuators have extensive spaceflight experience, most notably on the  Delta launch vehicle stage separation system. Pneumatic actuators possess larger specific force  capability (N/kg) than mechanical springs, giving 4 to 5 times the kickoff force of springs of the same  mass. Higher part count and pressuri zed components leads to pot entially lower reliability than the simpler  mechanical springs, making them less attractive for a crewed mission.   Gas thrusters: While compressed gas thrusters have aerospace flight heritage in solid rocket booster  separation (Figure 3) they have no known experience as a spacecraft or payload separation device to this  author’s knowledge. Their benefit is in producing a large specific force giving very high drive capability,  even greater than 10 times that of mechanical springs for the same mass. Where high kickoff forces are not required, their greater complexity and potentially lower reliability may make them less attractive. Gas  thrusters are currently being traded against pneumatic actuators for the Ares 1 launch vehicle staging  mechanism as well.  Figure 3. Pyrotechnic Gas Thruster/Actuator (left), Pneumatic Actuator System (right)   (Used with permission of Scot, Inc.)  Modeling Approach To address the mechanism design sizing, two modeling approaches were invoked. Each method allowed  for easy evaluations as vehicle configuration changes occurred. The first approach was a simplified  Simulink model to quickly assess the critical clearance between the Orion Engine nozzle and the SA. The  second approach involved the generation of an ADAMS 3D geometric model to assess secondary effects due to offset mass centers of gravity, off-diagonal moment of inertia terms and other out of plane effects.  It also captured any interference due to potential contact between other parts of the Orion Vehicle and the  SA. Simulink Model Approach The Simulink model approach [12] for the abort simulation was based on the translation and rotational  equations of motion, which are integrated through the time  step function to determine the relative  positions of the Ares Upper Stage and Orion. Figu re 4 depicts the full Simulink model. The separation  force is applied as either a constant pressure (as from a gas thruster) or a variable force (mechanical spring) over the length of the actuation. Assumptions of planar motion for the location of element centers  of gravity and constant component masses for the duration of the separation event are incorporated.  Capability has been added to the model to include the residual J-2X engine thrust acting on the Ares I US after separation and the contribution of Reaction Control System thrust to the separation acceleration.   348