Jamshid A. Samareh

An integrated tool for system analysis of sample return vehicles

The next important step in space exploration is the return of sample materials from extraterrestrial locations to Earth for analysis. Most mission concepts that return sample material to Earth share one common element: an Earth entry vehicle. The analysis and design of entry vehicles is multidisciplinary in nature, requiring the application of mass sizing, flight mechanics, aerodynamics, aerothermodynamics, thermal analysis, structural analysis, and impact analysis tools. Integration of a multidisciplinary problem is a challenging task; the execution process and data transfer among disciplines should be automated and consistent. This paper describes an integrated analysis tool for the design and sizing of an Earth entry vehicle. The current tool includes the following disciplines: mass sizing, flight mechanics, aerodynamics, aerothermodynamics, and impact analysis tools. Python and Java languages are used for integration. Results are presented and compared with the results from previous studies.

Human Mars lander design for NASA's evolvable mars campaign

Landing humans on Mars will require entry, descent, and landing capability beyond the current state of the art. Nearly twenty times more delivered payload and an order of magnitude improvement in precision landing capability will be necessary. To better assess entry, descent, and landing technology options and sensitivities to future human mission design variations, a series of design studies has been initiated. This paper describes the results of the first design study in the series of studies to be completed in 2016 and includes system and subsystem design details including mass and power estimates for a lander design using the Hypersonic Inflatable Aerodynamic Decelerator (HIAD) entry technology. Future design activities in this series will focus on other entry technology options.

Simulation and Analyses of Multi-Body Separation in Launch Vehicle Staging Environment

The development of methodologies, techniques, and tools for analysis and simulation of multi-body separation is critically needed for successful design and operation of next generation launch vehicles. As a part of this activity, ConSep simulation tool is being developed. ConSep is a generic MATLAB-based front-and-back-end to the commercially available ADAMS  solver, an industry standard package for solving multi-body dynamic problems. This paper discusses the 3-body separation capability in ConSep and its application to the separation of the Shuttle Solid Rocket Boosters (SRBs) from the External Tank (ET) and the Orbiter. The results are compared with STS-1 flight data.

Application of Constraint Force Equation Methodology for Launch Vehicle Stage Separation

A = axial location of the solid rocket booster (SRB) reference point in Space Shuttle main engine (SSME) plume coordinate system, ft ax, ay, az = acceleration components along body axes (excluding components due to gravity), ft=s CA = isolated (freestream) axial force coefficient Cm = isolated (freestream) pitching moment coefficient CN = isolated (freestream) normal force coefficient Cn = isolated (freestream) yawing moment coefficient CY = isolated (freestream) side force coefficient Cl = isolated (freestream) rolling moment coefficient eA, eB = unit vectors in body A and body B, respectively Fp = plume impingement force, lb Fx, Fy, Fz = aerodynamic forces in axial, lateral, and normal directions, lb F CON A , F CON B = joint constraint force vector for body A and body B

Systems analysis of carbon nanotubes: opportunities and challenges for space applications

Recent availability of carbon nanotubes (CNTs) in quantities and formats amenable to producing macroscale components invites consideration of these materials in space applications where their attractive properties can enable the realization of bold concepts for affordable space exploration. The challenge is to identify relevant systems and quantify the benefits at the systems level. Before significant investment or adoption of CNTs for large aerospace systems can be justified, there must be a plausible path to attain the perceived systems level benefits. This challenging step requires a close collaboration among experts on CNTs and aerospace system communities. This paper provides an overview of a few relevant potential CNTs applications for space systems and the gap that must be overcome for deployment of CNTs. It also provides a simple engineering-level systems analysis approach to quantify the benefits of using CNTs over state of the art material solutions.

Thermal Protection System Mass Estimating Relationships For Blunt-Body, Earth Entry Spacecraft

System analysis and design of any entry system must balance the level fidelity for each discipline against the project timeline. One way to inject high fidelity analysis earlier in the design effort is to develop surrogate models for the high-fidelity disciplines. Surrogate models for the Thermal Protection System (TPS) are formulated as Mass Estimating Relationships (MERs). The TPS MERs are presented that predict the amount of TPS necessary for safe Earth entry for blunt-body spacecraft using simple correlations that closely match estimates from NASA’s high-fidelity ablation modeling tool, the Fully Implicit Ablation and Thermal Analysis Program (FIAT). These MERs provide a first order estimate for rapid feasibility studies. There are 840 different trajectories considered in this study, and each TPS MER has a peak heating limit. MERs for the vehicle forebody include the ablators Phenolic Impregnated Carbon Ablator (PICA) and Carbon Phenolic atop Advanced Carbon-Carbon. For the aftbody, the materials are Silicone Impregnated Reusable Ceramic Ablator (SIRCA), Acusil II, SLA-561V, and LI-900. The MERs are accurate to within 14% (at one standard deviation) of FIAT prediction, and the most any MER under predicts FIAT TPS thickness is 18.7%. This work focuses on the development of these MERs, the resulting equations, model limitations, and model accuracy.

A Multifunctional Hot Structure Heat Shield Concept for Planetary Entry

A multifunctional hot structure heatshield concept is being developed to provide technology enhancements with significant benefits compared to the current state-of-the-art heatshield technology. These benefits can potentially enable future planetary missions. The concept is unique in integrating the function of the thermal protection system with the primary load carrying structural component. An advanced carbon-carbon material system has been evaluated for the load carrying structure, which will be utilized on the outer surface of the heatshield, and thus will operate as a hot structure exposed to the severe aerodynamic heating associated with planetary entry. Flexible, highly efficient blanket insulation is sized for use underneath the hot structure to maintain required operational internal temperatures. The approach followed includes developing preliminary designs to demonstrate feasibility of the concept and benefits over a traditional, baseline design. Where prior work focused on a concept for an Earth entry vehicle, the current efforts presented here are focused on developing a generic heatshield model and performing a trade study for a Mars entry application. This trade study includes both structural and thermal evaluation. The results indicate that a hot structure concept is a feasible alternative to traditional heatshields and may offer advantages that can enable future entry missions.

Preliminary Development of a Multifunctional Hot Structure Heat Shield

Development of a Multifunctional Hot Structure Heat Shield concept has initiated with the goal to provide advanced technology with significant benefits compared to the current state of the art heat shield technology. The concept is unique in integrating the function of the thermal protection system with the primary load carrying structural component. An advanced carbon-carbon material system has been evaluated for the load carrying structure, which will be utilized on the outer surface of the heat shield, and thus will operate as a hot structure exposed to the severe aerodynamic heating associated with planetary entry. Flexible, highly efficient blanket insulation has been sized for use underneath the hot structure to maintain desired internal temperatures. The approach was to develop a preliminary design to demonstrate feasibility of the concept. The preliminary results indicate that the concept has the potential to save both mass and volume with significantly less recession compared to traditional heat shield designs, and thus provide potential to enable new planetary missions.

Large Mass, Entry, Descent and Landing Sensitivity Results for Environmental, Performance, and Design Parameters

Landing on Mars has been a challenging task. Past NASA missions have shown resilience to increases in spacecraft mass by scaling back requirements such as landing site altitude, landing site location and arrival time. Knowledge of the partials relating requirements to mass is critical for mission designers to understand so that the project can retain margin throughout the process. Looking forward to new missions that will land 1.5 metric tons or greater, the current level of technology is insufficient, and new technologies will need to be developed. Understanding the sensitivity of these new technologies to requirements is the purpose of this paper.

Application of CFE/POST2 for Simulation of Launch Vehicle Stage Separation

The constraint force equation (CFE) methodology provides a framework for modeling constraint forces and moments acting at joints that connect multiple vehicles. With implementation in Program to Optimize Simulated Trajectories II (POST 2), the CFE provides a capability to simulate end-to-end trajectories of launch vehicles, including stage separation. In this paper, the CFE/POST2 methodology is applied to the Shuttle-SRB separation problem as a test and validation case. The CFE/POST2 results are compared with STS-1 flight test data.