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NASA NTRS2010

Spinoff 2010

Topics covered include: Burnishing Techniques Strengthen Hip Implants; Signal Processing Methods Monitor Cranial Pressure; Ultraviolet-Blocking Lenses Protect, Enhance Vision; Hyperspectral Systems Increase Imaging Capabilities; Programs Model the Future of Air Traffic Management; Tail Rotor Airfoils Stabilize Helicopters, Reduce Noise; Personal Aircraft Point to the Future of Transportation; Ducted Fan Designs Lead to Potential New Vehicles; Winglets Save Billions of Dollars in Fuel Costs; Sensor Systems Collect Critical Aerodynamics Data; Coatings Extend Life of Engines and Infrastructure; Radiometers Optimize Local Weather Prediction; Energy-Efficient Systems Eliminate Icing Danger for UAVs; Rocket-Powered Parachutes Rescue Entire Planes; Technologies Advance UAVs for Science, Military; Inflatable Antennas Support Emergency Communication; Smart Sensors Assess Structural Health; Hand-Held Devices Detect Explosives and Chemical Agents; Terahertz Tools Advance Imaging for Security, Industry; LED Systems Target Plant Growth; Aerogels Insulate Against Extreme Temperatures; Image Sensors Enhance Camera Technologies; Lightweight Material Patches Allow for Quick Repairs; Nanomaterials Transform Hairstyling Tools; Do-It-Yourself Additives Recharge Auto Air Conditioning; Systems Analyze Water Quality in Real Time; Compact Radiometers Expand Climate Knowledge; Energy Servers Deliver Clean, Affordable Power; Solutions Remediate Contaminated Groundwater; Bacteria Provide Cleanup of Oil Spills, Wastewater; Reflective Coatings Protect People and Animals; Innovative Techniques Simplify Vibration Analysis; Modeling Tools Predict Flow in Fluid Dynamics; Verification Tools Secure Online Shopping, Banking; Toolsets Maintain Health of Complex Systems; Framework Resources Multiply Computing Power; Tools Automate Spacecraft Testing, Operation; GPS Software Packages Deliver Positioning Solutions; Solid-State Recorders Enhance Scientific Data Collection; Computer Models Simulate Fine Particle Dispersion; Composite Sandwich Technologies Lighten Components; Cameras Reveal Elements in the Short Wave Infrared; Deformable Mirrors Correct Optical Distortions; Stitching Techniques Advance Optics Manufacturing; Compact, Robust Chips Integrate Optical Functions; Fuel Cell Stations Automate Processes, Catalyst Testing; Onboard Systems Record Unique Videos of Space Missions; Space Research Results Purify Semiconductor Materials; and Toolkits Control Motion of Complex Robotics.

Source record↗
NASA NTRS1993

Automated propellant leak detection

An automated hydrogen leak detection system is being developed for earth-to-orbit rocket engine applications. The system consists of three elements, a sensor array, a signal processing unit, and a diagnostic processor. The sensor array consists of discrete solid state sensors which are located at specific potential leak sites and in potential leak zones. The signal processing unit provides excitations power for the sensors and provides analog to digital data conversion of the sensor signals. The diagnostic computer analyzes the sensor outputs to determine leak sources and magnitude. Leak data from the sensor network is interpreted using knowledge based software and displayed on-line through a graphical user interface including 3-D leak visualization. The system requirements have been developed assuming eventual application to the Space Shuttle Main Engine which requires approximately 72 measurement locations. A prototype system has been constructed to demonstrate the operational features of the system. This system includes both prototype electronics and data processing software algorithms. Experiments are in progress to evaluate system operation at conditions which simulate prelaunch and flight. The prototype system consists of a network using 16 sensors within a testbed which simulates engine components. Sensor response, orientation, and data analysis algorithms are being evaluated using calibrated leaks produced within the testbed. A prototype flight system is also under development consisting of 8 sensors and flight capable electronics with autonomous control and data recording.

Makel, D. B.↗
NASA NTRS1994

Assessment of impact damage of composite rocket motor cases

This contract reviewed the available literature on mechanisms of low velocity impact damage in filament wound rocket motor cases, MDE methods to quantify damage, critical coupon level test methods, manufacturing and material process variables and empirical and analytical modeling off impact damage. The critical design properties for rocket motor cases are biaxial hoop and axial tensile strength. Low velocity impact damage is insidious because it can create serious nonvisible damage at very low impact velocities. In thick rocket motor cases the prevalent low velocity impact damage is fiber fracture and matrix cracking adjacent to the front face. In contrast, low velocity loading of thin wall cylinders induces flexure, depending on span length and the flexure induces delamination and tensile cracking on the back face wall opposed to impact occurs due to flexural stresses imposed by impact loading. Important NDE methods for rocket motor cases are non-contacting methods that allow inspection from one side. Among these are vibrothermography, and pulse-echo methods based on acoustic-ultrasonic methods. High resolution techniques such as x-ray computed tomography appear to have merit for accurate geometrical characterization of local damage to support development of analytical models of micromechanics. The challenge of coupon level testing is to reproduce the biaxial stress state that the full scale article experiences, and to determine how to scale the composite structure to model full sized behavior. Biaxial tensile testing has been performed by uniaxially tensile loading internally pressurized cylinders. This is experimentally difficult due to gripping problems and pressure containment. Much prior work focused on uniaxial tensile testing of model filament wound cylinders. Interpretation of the results of some studies is complicated by the fact that the fabrication process did not duplicate full scale manufacturing. It is difficult to scale results from testing subscale cylinders since there are significant differences in out time of the resins relative to full scale cylinder fabrication, differences in hoop fiber tensioning and unsatisfactory coupon configurations. It appears that development of a new test method for subscale cylinders is merited. Damage tolerance may be improved by material optimization that uses fiber treatments and matrix modifications to control the fiber matrix interface bonding. It is difficult to develop process optimization in subscale cylinders without also modeling the longer out times resins experience in full scale testing. A major breakthrough in characterizing the effect of impact damage on residual strength, and understanding how to scale results of subscale evaluations, will be a sound micromechanical model that described progressive failure of the composite. Such models will utilize a three dimensional stress analysis due to the complex nature of low velocity impact stresses in thick composites. When these models are coupled with non-contact NDE methods that geometrically characterize the damage and acoustic methods that characterize the effective local elastic properties, accurate assessment of residual strength from impact damage may be possible. Directions for further development are suggested.

Paris, Henry G.↗
NASA NTRS2007

Did Vertigo Kill America's Forgotten Astronaut?

On November 15, 1967, U.S. Air Force test pilot Major Michael J. Adams was killed while flying the X-15 rocket-propelled research vehicle in a parabolic spaceflight profile. This flight was part of a joint effort with NASA. An electrical short in one of the experiments aboard the vehicle caused electrical transients, resulting in excessive workload by the pilot. At altitude Major Adams inappropriately initiated a flat spin that led to a series of unusual aircraft attitudes upon atmospheric re-entry, ultimately causing structural failure of the airframe. Major Adams was known to experience vertigo (i.e. spatial disorientation) while flying the X-15, but all X-15 pilots most likely experienced vertigo (i.e. somatogravic, or "Pitch-Up", illusion) as a normal physiologic response to the accelerative forces involved. Major Adams probably experienced vertigo to a greater degree than did others, since prior aeromedical testing for astronaut selection at Brooks AFB revealed that he had an unusually high degree of labyrinthine sensitivity. Subsequent analysis reveals that after engine burnout, and through the zenith of the flight profile, he likely experienced the oculoagravic ("Elevator") illusion. Nonetheless, painstaking investigation after the mishap revealed that spatial disorientation (Type II, Recognized) was NOT the cause, but rather, a contributing factor. The cause was in fact the misinterpretation of a dual-use flight instrument (i.e. Loss of Mode Awareness), resulting in confusion between yaw and roll indications, with subsequent flight control input that was inappropriate. Because of the altitude achieved on this flight, Major Adams was awarded Astronaut wings posthumously. Understanding the potential for spatial disorientation, particularly the oculoagravic illusion, associated with parabolic spaceflight profiles, and understanding the importance of maintaining mode awareness in the context of automated cockpit design, are two lessons that have direct application to the commercial space industry today.

Bendrick, Gregg A.↗
NASA NTRSDate not supplied

Reducing Landing Site Contamination Using 3-D Trajectory Optimization for Surface Hoppers

Rocket-powered vehicles utilizing Vertical Take-off Vertical Landing (VTVL) are a compelling alternative to surface rovers for exploring planetary and lunar bodies. These so called “hoppers” provide enhanced mobility for accessing locations difficult to reach, and over a wider region of the surface. However, contamination and plume interactions from rocket exhaust deposited at landing sites is anticipated since landing approaches are typically along a vertical direction during the final descent. Consequently, exhaust products may alter the surface chemistry, potentially confounding compositional analysis for samples collected in the vicinity of the landing site or jeopardize mining efforts. There has been no rigorous study on flight maneuvers that can mitigate plume-to-surface interactions. A multi-objective optimization tool has been developed to simulate propulsive hops on a planetary body and minimize both fuel consumption and site alterations. Trajectories are derived by multi-objective optimization and include solutions with significant reduction in contamination for a modest increase in fuel consumption. For these solutions, surface-to-surface propulsive transfer is demonstrated, but the method can also be modified for orbit-to-surface transfers (e.g., landers).

guidance↗
NASA NTRSDate not supplied

Reducing Landing Site Contamination Using 3-D Trajectory Optimization for Surface Hoppers

Rocket-powered vehicles utilizing Vertical Take-off Vertical Landing (VTVL) are a compelling alternative to surface rovers for exploring planetary and lunar bodies. These so called “hoppers” provide enhanced mobility for accessing locations difficult to reach, and over a wider region of the surface. However, contamination and plume interactions from rocket exhaust deposited at landing sites is anticipated since landing approaches are typically along a vertical direction during the final descent. Consequently, exhaust products may alter the surface chemistry, potentially confounding compositional analysis for samples collected in the vicinity of the landing site or jeopardize mining efforts. There has been no rigorous study on flight maneuvers that can mitigate plume-to-surface interactions. A multi-objective optimization tool has been developed to simulate propulsive hops on a planetary body and minimize both fuel consumption and site alterations. Trajectories are derived by multi-objective optimization and include solutions with significant reduction in contamination for a modest in-crease in fuel consumption. For these solutions, surface-to-surface propulsive transfer is demonstrated, but the method can also be modified for orbit-to-surface transfers (e.g., landers).

guidance↗
NASA NTRS2003

Propulsion Flight-Test Fixture

NASA Dryden Flight Research Center s new Propulsion Flight Test Fixture (PFTF), designed in house, is an airborne engine-testing facility that enables engineers to gather flight data on small experimental engines. Without the PFTF, it would be necessary to obtain such data from traditional wind tunnels, ground test stands, or laboratory test rigs. Traditionally, flight testing is reserved for the last phase of engine development. Generally, engines that embody new propulsion concepts are not put into flight environments until their designs are mature: in such cases, either vehicles are designed around the engines or else the engines are mounted in or on missiles. However, a captive carry capability of the PFTF makes it possible to test engines that feature air-breathing designs (for example, designs based on the rocket-based combined cycle) economically in subscale experiments. The discovery of unknowns made evident through flight tests provides valuable information to engine designers early in development, before key design decisions are made, thereby potentially affording large benefits in the long term. This is especially true in the transonic region of flight (from mach 0.9 to around 1.2), where it can be difficult to obtain data from wind tunnels and computational fluid dynamics. In January 2002, flight-envelope expansion to verify the design and capabilities of the PFTF was completed. The PFTF was flown on a specially equipped supersonic F-15B research testbed airplane, mounted on the airplane at a center-line attachment fixture, as shown in Figure 1. NASA s F-15B testbed has been used for several years as a flight-research platform. Equipped with extensive research air-data, video, and other instrumentation systems, the airplane carries externally mounted test articles. Traditionally, the majority of test articles flown have been mounted at the centerline tank-attachment fixture, which is a hard-point (essentially, a standardized weapon-mounting fixture). This hard-point has large weight margins, and, because it is located near the center of gravity of the airplane, the weight of equipment mounted there exerts a minimal effect on the stability and controllability of the airplane. The PFTF (see Figure 2) includes a one-piece aluminum structure that contains space for instrumentation, propellant tanks, and feed-system components. The PFTF also houses a force balance, on which is mounted the subscale engine or other experimental apparatus that is to be the subject of a flight test. The force balance measures a combination of inertial and aerodynamic forces and moments acting on the experimental apparatus.

Palumbo, Nate↗
NASA NTRS1994

STS-61 Space Shuttle mission report

The STS-61 Space Shuttle Program Mission Report summarizes the Hubble Space Telescope (HST) servicing mission as well as the Orbiter, External Tank (ET), Solid Rocket Booster (SRB), Redesigned Solid Rocket Motor (RSRM), and the Space Shuttle main engine (SSME) systems performance during the fifty-ninth flight of the Space Shuttle Program and fifth flight of the Orbiter vehicle Endeavour (OV-105). In addition to the Orbiter, the flight vehicle consisted of an ET designated as ET-60; three SSME's which were designated as serial numbers 2019, 2033, and 2017 in positions 1, 2, and 3, respectively; and two SRB's which were designated BI-063. The RSRM's that were installed in each SRB were designated as 360L023A (lightweight) for the left SRB, and 360L023B (lightweight) for the right SRB. This STS-61 Space Shuttle Program Mission Report fulfills the Space Shuttle Program requirement as documented in NSTS 07700, Volume 8, Appendix E. That document requires that each major organizational element supporting the Program report the results of its hardware evaluation and mission performance plus identify all related in-flight anomalies. The primary objective of the STS-61 mission was to perform the first on-orbit servicing of the Hubble Space Telescope. The servicing tasks included the installation of new solar arrays, replacement of the Wide Field/Planetary Camera I (WF/PC I) with WF/PC II, replacement of the High Speed Photometer (HSP) with the Corrective Optics Space Telescope Axial Replacement (COSTAR), replacement of rate sensing units (RSU's) and electronic control units (ECU's), installation of new magnetic sensing systems and fuse plugs, and the repair of the Goddard High Resolution Spectrometer (GHRS). Secondary objectives were to perform the requirements of the IMAX Cargo Bay Camera (ICBC), the IMAX Camera, and the Air Force Maui Optical Site (AMOS) Calibration Test.

Fricke, Robert W., Jr.↗
NASA NTRS2009

Recovering Aerodynamic Side Loads on Rocket Nozzles using Quasi-Static Strain-Gage Measurements

During over-expanded operation of rocket nozzles, which is defined to be when the exit pressure is greater than internal pressure over some part of the nozzle, the nozzle will experience a transverse forcing function due to the pressure differential across the nozzle wall. Over-expansion occurs during the nozzle start-up and shutdown transient, even in high-altitude engines, because most test facilities cannot completely reproduce the near-vacuum pressures at those altitudes. During this transient, the pressure differential moves axially down the nozzle as it becomes pressurized, but this differential is never perfectly symmetric circumferentially. The character of the forcing function is highly complex and defined by a series of restricted and free shock separations. The subject of this paper is the determination of the magnitude of this loading during sub-scale testing via measurement of the structural dynamic response of the nozzle and its support structure. An initial attempt at back-calculating this load using the inverse of the transfer function was performed, but this attempt was shown to be highly susceptible to numerical error. The final method chosen was to use statically calibrated strain data and to filter out the system fundamental frequency such that the measured response yields close to the correct dynamic loading function. This method was shown to capture 93% of the pressure spectral energy using controlled load shaker testing. This method is one of the only practical ways for the inverse determination of the forcing function for non-stationary excitations, and, to the authors' knowledge, has not been described in the literature to date.

Brown, Andrew↗
NASA NTRS2011

The Dawn of Vesta Science

The Dawn mission is part of NASA's Discovery Program and has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total ?V of 11.3 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Fullpower thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional ?V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. An additional 210 hours of IPS thrusting was used to enter the first Vesta science orbit, called Survey orbit, on August 3, 2011 at an altitude of approximately 2,735 km. To date the IPS has been operated for 23,621 hours, consumed approximately 252 kg of xenon, and provided a delta-V of approximately 6.7 km/s. The IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. The only significant issue in over the almost four years of IPS operations in flight was the temporary failure of a valve driver board in the Digital Control and Interface Unit-1 (DCIU-1), resulting in a loss of thrust of approximately 29 hours. Thrusting operations resumed after switching to DCIU-2, and power cycling conducted after orbit capture restored DCIU-1 to full functionality. After about three weeks of Survey orbit operations the IPS will be used to transfer the spacecraft to the other planned science orbit altitudes. After approximately one year of science operations the IPS will be used for escape from Vesta and then for thrusting to Ceres with a planned arrival date at Ceres in February 2015. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations through the start of science operations at Vesta.

electric propulsion↗
NASA NTRS2010

Monitoring System for Storm Readiness and Recovery of Test Facilities: Integrated System Health Management (ISHM) Approach

Severe weather events are likely occurrences on the Mississippi Gulf Coast. It is important to rapidly diagnose and mitigate the effects of storms on Stennis Space Center's rocket engine test complex to avoid delays to critical test article programs, reduce costs, and maintain safety. An Integrated Systems Health Management (ISHM) approach and technologies are employed to integrate environmental (weather) monitoring, structural modeling, and the suite of available facility instrumentation to provide information for readiness before storms, rapid initial damage assessment to guide mitigation planning, and then support on-going assurance as repairs are effected and finally support recertification. The system is denominated Katrina Storm Monitoring System (KStorMS). Integrated Systems Health Management (ISHM) describes a comprehensive set of capabilities that provide insight into the behavior the health of a system. Knowing the status of a system allows decision makers to effectively plan and execute their mission. For example, early insight into component degradation and impending failures provides more time to develop work around strategies and more effectively plan for maintenance. Failures of system elements generally occur over time. Information extracted from sensor data, combined with system-wide knowledge bases and methods for information extraction and fusion, inference, and decision making, can be used to detect incipient failures. If failures do occur, it is critical to detect and isolate them, and suggest an appropriate course of action. ISHM enables determining the condition (health) of every element in a complex system-of-systems or SoS (detect anomalies, diagnose causes, predict future anomalies), and provide data, information, and knowledge (DIaK) to control systems for safe and effective operation. ISHM capability is achieved by using a wide range of technologies that enable anomaly detection, diagnostics, prognostics, and advise for control: (1) anomaly detection algorithms and strategies, (2) fusion of DIaK for anomaly detection (model-based, numerical, statistical, empirical, expert-based, qualitative, etc.), (3) diagnostics/prognostics strategies and methods, (4) user interface, (5) advanced control strategies, (6) integration architectures/frameworks, (7) embedding of intelligence. Many of these technologies are mature, and they are being used in the KStorMS. The paper will describe the design, implementation, and operation of the KStorMS; and discuss further evolution to support other needs such as condition-based maintenance (CBM).

Figueroa, Fernando↗
NASA NTRSDate not supplied

Space Transformation -- Localizing the Remote and Connecting the Isolated

In motivating the Space Transformation theme for this year’s 4S symposium, the organizers provided the following context, “Transformation of economies are driven by a change in values and accelerated by new technologies.” These words rang particularly true when I read them at the beginning of the holiday season. Like so many others, I was in the early phases of my Christmas shopping procrastination campaign, and I’d just been reflecting on how Amazon Prime was the transformational tool I’d been waiting for. Basic limiting principles of time and space, supply and demand, were all but erased by the Amazon Prime phenomenon. Coupled with emerging 3D printing and other adaptive manufacturing technologies, a transformation from deliberate planning to “think it … have it” had occurred, empowering me to procrastinate longer than I’d ever dreamed possible. The organizers went on to ponder, “Will space transformation also affect society?”, just as our team at the Air Force Research Lab’s (AFRL) Center for Rapid Innovation (CRI) were working alongside partners within our larger Integrated Capabilities Directorate, NASA’s Flight Opportunities and Small Spacecraft Technology programs, and DARPA’s Luna-10 program to develop technologies and execute demonstration missions that leverage the space domain to genuinely connect even the most remote and austere domains on the timeline of need. Picking apart the miracle that is Amazon prime, where does the model fail, and why? More relevantly to the theme of this year’s symposium, how can the space domain be used to overcome its limitations and minimize its weaknesses? Perhaps it is best assessed in the context of Use Cases. What are the Amazon delivery cost, schedule, and cargo limiters to the Amundsen-Scott South Pole Research Station, or the Lunar South Pole Research Station? This paper will explore enabling infrastructure that allows Amazon prime to thrive and assess the transformational enabling technologies that would be necessary to extend that miracle to the truly remote or the truly austere. Localizing the Remote • First, it will evaluate the ability of the on-going AFRL Rocket Cargo and Space Initiatives Ringside Seats systems, coupled with Astrobotic’s Xodiak and Xogdor capabilities, developed to support the NASA Flight Opportunities Program (FOP), to supply orbital/suborbital delivery to both improved and austere sites on the Earth and Moon. • Then, it will add the surface terminal distribution leg, with an examination of Lunar Outpost’s Mobile Autonomous Prospecting Platform (MAPP), equipped with Mobile Autonomous Robotic Swarm (MARS) software, and Intuitive Machine’s Hopper, developed with support of AFRL and NASA’s Commercial Lunar Payload Services (CLPS) program. Connecting the Isolated From there, it will focus on the destination, asking what implied destination services are required to support highly assured autonomous delivery. • Specifically, it will highlight Astrobotic’s Skymage mesh-networked publish and subscribe communication and navigation service, as well as AFRL’s on-going developments of radioisotope and reactor nuclear-sourced thermoelectric power generation and distribution systems under development under the Joint Emergent Technology Supplying On-orbit Nuclear Power (JETSON) program by Lockheed Martin, Westinghouse, Intuitive Machines, and Zeno Power, to provide the power service to locations well off the grid. • Finally, the paper will connect to the “human machine”. What connects the remote or in-situ human consumer to the remote domain? What connects the diverse international government and commercial services to each other? The former will focus on AFRL’s OraCloud feeding their Space Defense Control and Characterization System (SDCCS) and Lunar Station’s MoonHacker systems, while the latter will focus on the BlueHalo/Tensor LunX Technology Platform for the Cislunar Commodity Marketplace. In 1984, Krafft Ehricke famously remarked that, “If God wanted man to become a spacefaring species, He would have given man a Moon.” This paper is not about the Moon, but is about humans as a spacefaring species, shedding the pesky land/air limitations of the Amazon Prime model … so that we can all live a procrastinator’s “think it … have it” existence.

Charles Finley↗
NASA NTRS2009

Modeling and Simulation of Variable Mass, Flexible Structures

The advent of the new Ares I launch vehicle has highlighted the need for advanced dynamic analysis tools for variable mass, flexible structures. This system is composed of interconnected flexible stages or components undergoing rapid mass depletion through the consumption of solid or liquid propellant. In addition to large rigid body configuration changes, the system simultaneously experiences elastic deformations. In most applications, the elastic deformations are compatible with linear strain-displacement relationships and are typically modeled using the assumed modes technique. The deformation of the system is approximated through the linear combination of the products of spatial shape functions and generalized time coordinates. Spatial shape functions are traditionally composed of normal mode shapes of the system or even constraint modes and static deformations derived from finite element models of the system. Equations of motion for systems undergoing coupled large rigid body motion and elastic deformation have previously been derived through a number of techniques [1]. However, in these derivations, the mode shapes or spatial shape functions of the system components were considered constant. But with the Ares I vehicle, the structural characteristics of the system are changing with the mass of the system. Previous approaches to solving this problem involve periodic updates to the spatial shape functions or interpolation between shape functions based on system mass or elapsed mission time. These solutions often introduce misleading or even unstable numerical transients into the system. Plus, interpolation on a shape function is not intuitive. This paper presents an approach in which the shape functions are held constant and operate on the changing mass and stiffness matrices of the vehicle components. Each vehicle stage or component finite element model is broken into dry structure and propellant models. A library of propellant models is used to describe the distribution of mass in the fuel tank or Solid Rocket Booster (SRB) case for various propellant levels. Based on the mass consumed by the liquid engine or SRB, the appropriate propellant model is coupled with the dry structure model for the stage. Then using vehicle configuration data, the integrated vehicle model is assembled and operated on by the constant system shape functions. The system mode shapes and frequencies can then be computed from the resulting generalized mass and stiffness matrices for that mass configuration. The rigid body mass properties of the vehicle are derived from the integrated vehicle model. The coupling terms between the vehicle rigid body motion and elastic deformation are also updated from the constant system shape functions and the integrated vehicle model. This approach was first used to analyze variable mass spinning beams and then prototyped into a generic dynamics simulation engine. The resulting code was tested against Crew Launch Vehicle (CLV-)class problems worked in the TREETOPS simulation package and by Wilson [2]. The Ares I System Integration Laboratory (SIL) is currently being developed at the Marshall Space Flight Center (MSFC) to test vehicle avionics hardware and software in a hardware-in-the-loop (HWIL) environment and certify that the integrated system is prepared for flight. The Ares I SIL utilizes the Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) tool to simulate the launch vehicle and stimulate avionics hardware. Due to the presence of vehicle control system filters and the thrust oscillation suppression system, which are tuned to the structural characteristics of the vehicle, ARTEMIS must incorporate accurate structural models of the Ares I launch vehicle. The ARTEMIS core dynamics simulation models the highly coupled nature of the vehicle flexible body dynamics, propellant slosh, and vehicle nozzle inertia effects combined with mass and flexible body properties that vary significant with time during the flight. All forces that act on the vehicle during flight must be simulated, including deflected engine thrust force, spatially distributed aerodynamic forces, gravity, and reaction control jet thrust forces. These forces are used to excite an integrated flexible vehicle, slosh, and nozzle dynamics model for the vehicle stack that simulates large rigid body translations and rotations along with small elastic deformations. Highly effective matrix math operations on a distributed, threaded high-performance simulation node allow ARTEMIS to retain up to 30 modes of flex for real-time simulation. Stage elements that separate from the stack during flight are propagated as independent rigid six degrees of freedom (6DOF) bodies. This paper will present the formulation of the resulting equations of motion, solutions to example problems, and describe the resulting dynamics simulation engine within ARTEMIS.

Tobbe, Patrick A.↗
NASA NTRS2015

Printing Outside the Box: Additive Manufacturing Processes for Fabrication of Large Aerospace Structures

To achieve NASA's mission of space exploration, innovative manufacturing processes are being applied to the fabrication of propulsion elements. Liquid rocket engines (LREs) are comprised of a thrust chamber and nozzle extension as illustrated in figure 1 for the J2X upper stage engine. Development of the J2X engine, designed for the Ares I launch vehicle, is currently being incorporated on the Space Launch System. A nozzle extension is attached to the combustion chamber to obtain the expansion ratio needed to increase specific impulse. If the nozzle extension could be printed as one piece using free-form additive manufacturing (AM) processes, rather than the current method of forming welded parts, a considerable time savings could be realized. Not only would this provide a more homogenous microstructure than a welded structure, but could also greatly shorten the overall fabrication time. The main objective of this study is to fabricate test specimens using a pulsed arc source and solid wire as shown in figure 2. The mechanical properties of these specimens will be compared with those fabricated using the powder bed, selective laser melting technology at NASA Marshall Space Flight Center. As printed components become larger, maintaining a constant temperature during the build process becomes critical. This predictive capability will require modeling of the moving heat source as illustrated in figure 3. Predictive understanding of the heat profile will allow a constant temperature to be maintained as a function of height from substrate while printing complex shapes. In addition, to avoid slumping, this will also allow better control of the microstructural development and hence the properties. Figure 4 shows a preliminary comparison of the mechanical properties obtained.

Babai, Majid↗
NASA NTRS2025

Photophoretic Propulsion Enabling Mesosphere Exploration NIAC Phase I Final Report

This Phase I report presents a comprehensive study on photophoretic flyers—innovative, ultralight, solar-powered vehicles that harness photophoretic forces generated via Knudsen pumping to achieve sustained flight in the mesosphere (50–80 km altitude). By integrating advanced materials such as nanocardboard— characterized by its extremely low areal density (~1 g/m²) and high bending stiffness—with ultrathin light-absorbing coatings, the project converts incident solar radiation directly into a directed thrust. Extensive experimental investigations, coupled with high-fidelity computational fluid dynamics (CFD) simulations using ANSYS Fluent, validate the concept across various three-dimensional geometries, including spherical, conical, and rocket-shaped configurations. These simulations bridge the gap between free-molecular and continuum flow regimes, demonstrating that optimized designs can generate lift forces sufficient to support kilogram-scale payloads even in low-pressure environments. At the heart of this innovation is the use of Knudsen pumping, where temperature gradients across porous surfaces induce directional gas flow, creating a modest overpressure that provides lift. The report introduces an analytical framework that interpolates between the well-known low-Reynolds number drag regime and the high-Reynolds number momentum theory. This model accurately predicts lift based on design parameters such as microchannel dimensions, porous wall geometry, areal density, and nozzle exit area. For instance, simulations indicate that 10-meter-scale structures with carefully engineered porous walls can achieve the necessary pressure differential to support scientifically significant payloads (~1 kg). The study also explores a hybrid propulsion approach that combines solar buoyancy with photophoretic lift. Initially, solar heating creates a buoyant force that elevates the flyer to mesospheric altitudes. Once in the optimal pressure range, the photophoretic mechanism—powered by Knudsen pumping—takes over as the primary source of lift, ensuring stable, long-duration flight. This dual-mode operation not only facilitates the deployment of photophoretic flyers but also broadens the potential applications for mesospheric exploration. In addition to propulsion, the report investigates the integration of photophoretic thrusters for trajectory control of existing research balloons in the upper stratosphere. By dynamically adjusting the nozzle orientation and controlling the flow-through velocity, these thrusters provide precise maneuverability, enabling the flyers to counteract atmospheric disturbances and adjust their flight paths in real time. For example, a photophoretic thruster approximately 7.5 by 7.5 meters in size could be unfolded below a payload gondola of a 60 million-cubic-foot zero-pressure balloon. Such a thruster can provide horizontal speed control of approximately 1 m/s using only sunlight and no moving parts (except those needed to track the Sun and control the jet direction). Importantly, photophoretic thrusters operate more efficiently at higher altitudes, which is complementary to known trajectory control techniques, such as propellers and tethered wings, which are more effective at lower altitudes. Finally, the report identifies three scientific research thrusts where mesospheric aircraft technology can have a profound impact: atmospheric tides, characterization of gravity waves, and investigation of mesospheric instabilities. Overall, the findings of this Phase I project represent a significant advancement in photophoretic propulsion technology. By demonstrating that large-scale, ultralight structures can be powered solely by solar radiation—via carefully engineered Knudsen pumping—this work lays a robust foundation for scalable, near-space flight architectures. Future refinements in material fabrication, structural optimization, and integrated trajectory control are expected to further enhance performance, paving the way for operational demonstrations that could revolutionize atmospheric science, remote sensing, and communication networks.

Knudsen Pump↗
NASA NTRS2006

Space Flight 101

This viewgraph presentation reviews many aspects of spaceflight. There are many pictures of the International Space Station. Some of the topics covered in this review are: Have you ever wondered why we have launch windows? Or why the attitude of the Space Station changes? The half-day seminar answers some of the many questions about why and how we fly in space. Topics in the course were: What's so valuable about micro gravity? How do we get to micro G so close to a huge gravity well like the Earth? How come such a big rocket gets so little payload to space? Why do we have daily launch windows, and why are they so short? What's the beta angle, why does it change so strangely, and why do we care so much about it? Why do we have launch seasons for the Shuttle? Why can't we just launch any old day? Why do we see the station some days, not on others, and at different times and directions? Why do we keep changing the attitude of the Space Station? What are the certified attitudes of the station, and why did we pick these few? Why do we keep changing the altitude of the Space Station among these three? What's the difference between Power Balance, Energy Balance, and Depth-of-Discharge? Where does all the uncertainty come from in our orbit predictions for phasing, collisions, communications coverage, etc? Why do we usually reboost only on days that we do attitude changes? What's F(10.7), what does it do, why do we care, and why does it vary so much? Why do we care about orbital phasing of the ISS? Can't we just do phasing with the arriving vehicles? Why is the Space Station built the way it is? What's Sun Slicer? (or Night Glider, or Dual Angle, or (coming soon!) Mixmaster, or Outrigger, or...) What's a BGA, and what is BGA conditioning all about? What's a Control Moment Gyroscope, and what does it do? What's a desat? Why is it more trouble now than it used to be? How much orbital debris is there, and how dangerous is it? Why aren't we more worried about meteor storms? What is atomic Oxygen? Where does it come from, and why is it a nuisance? What's a gravity gradient attitude, and why don't we use it? How come we see things like phantom torques and phantom spikes and other things that Shuttles and stations and capsules have never seen? and How are all these topics affecting our assembly plans for this year and through the life of station?

Bacon, Jack↗
NASA NTRS2007

The X-43A Flight Research Program: Lessons Learned on the Road to Mach 10

During an aerospace engineer's undergraduate studies, he or she will attend classes in aerodynamics, thermodynamics, structures, stability and control, dynamics, design, propulsion, and computer science, along with the related courses in mathematics, physics, statistics, and chemistry required to understand the material. Upon graduation, the new engineer will have acquired a basic knowledge of how to build an aerospace vehicle. What only comes through experience, however, is the understanding of the inevitable imperfect process through which an aerospace vehicle is built. This is the adventure of turning a basic concept into functional hardware. Engineers working on a project must often deal with ambiguous situations. They are routinely asked by management to provide risk assessments of a project, yet even after careful analysis uncertainties remain. The project must be accomplished within finite limits of time and money. The question an engineer answers is whether the solution to potential problem is worth the cost and schedule delay, or if the solution might actually be worse than the problem it is meant to solve. Review protocols are established to ensure that an unknown has not been overlooked. But these cannot protect against an unknown unknown. Examples of these situations can be found in the history of the X-43A Hyper-X (Hypersonic Experiment) program. In this NASA project, a supersonic combustion ramjet (scramjet) engine was flight tested on a subscale vehicle. The X-43A Hyper-X Research Vehicle (HXRV) was launched from a B-52B mothership, then boosted to the test speed by a modified Pegasus rocket first stage, called the Hyper-X Launch Vehicle (HXLV). Once at the proper speed and altitude, the X-43A separated from the booster, stabilized itself, and then the engine test began. Although wind-tunnel scramjet engine tests had begun in the late 1950s, before the Hyper-X program there had never been an actual in-flight test of such an engine integrated with an appropriate airframe. Thus, while the scramjet had successfully operated in the artificial airflow of wind tunnels, the concept had yet to be proven in real air. These conditions meant changes in density and temperature, as well as changes in angle of attack and sideslip of a free-flying vehicle. A wind tunnel is limited in its ability to simulate these subtle factures, which have a major impact on almost any vehicle, but especially that of a scramjet's performance. The Hyper-X project was to provide a real-world benchmark of the ground test data. The full scale X-43A engine would be operated in the wind tunnel, and then flown, and the data from its operation would then be compared with projections. If these matched, the wind tunnel data would be considered a reliable design tool for future scramjet. If there were significant differences, the reasons for these would have to be identified. Until such information was available, scramjets would lack the technological maturity to be considered for future space launch or high-speed atmospheric flight vehicles.

Peebles, Curtis↗
NASA NTRS2017

Elastomeric Thermal Insulation Design Considerations in Long, Aluminized Solid Rocket Motors

An all-new sounding rocket was designed at NASA's Marshall Space Flight Center that featured an aft finocyl, aluminized solid propellant grain and silica-filled ethylene-propylene-diene monomer (SFEPDM) internal insulation. Upon the initial static firing of the first of this new design, the solid rocket motor (SRM) case failed thermally just upstream of the aft closure early in the burn time. Subsequent fluid modeling indicated that the high-velocity combustion-product jets emanating from the fin-slots in the propellant grain were likely inducing a strongly swirling flow, thus substantially increasing the severity of the convective environment on the exposed portion of the SFEPDM insulation in this region. The aft portion of the fin-slots in another of the motors were filled with propellant to eliminate the possibility of both direct jet impingement on the exposed SFEPDM and the appearance of strongly swirling flow in the aft region of the motor. When static-fired, this motor's case still failed in the same axial location, and, though somewhat later than for the first static firing, still in less than 1/3rd of the desired burn duration. These results indicate that the extreme material decomposition rates of the SFEPDM in this application are not due to gas-phase convection or shear but rather to interactions with burning aluminum or alumina slag. Further comparisons with between SFEPDM performance in this design and that in other hot-fire tests provide insight into the mechanisms of SFEPDM decomposition in SRM aft domes that can guide the upcoming redesign effort, as well as other future SRM designs. These data also highlight the current limitations of modeling elastomeric insulators solely with diffusion-controlled, gas-phase thermochemistry in SRM regions with significant viscous shear and/or condense-phase impingement or flow.

Martin, Heath T.↗