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At least 523 records · Page 29

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NASA NTRSDate not supplied

Blade Displacement Measurements of a Rotor in Forward Flight in the Langley 14- by 22-Ft Wind Tunnel

Stereo photogrammetry was used to measure the elastic bending and twist of a three-bladed rotor in forward flight in the NASA Langley 14-by 22-Ft Subsonic Wind Tunnel. The rotor was imaged from below by two cameras mounted at windows of a large turntable in the floor of the test section. At each test condition, the turntable and cameras were rotated and the cameras were triggered to allow measurements over a range of blade azimuth angles. Retro-reflective targets were applied along the leading and trailing edges of the lower surface of each blade. Image blur due to blade motion was minimized by illuminating the targets with high-intensity LED flash lamps mounted next to each camera. The cameras were re-calibrated at each angular position of the turntable using the positions in each image of retro-reflective targets on the ceiling of the test section whose spatial coordinates had been very accurately measured. The blade displacement measurements yielded elastic bending and twist along each blade as functions of blade azimuth, thrust coefficient, and advance ratio, all at a constant rotor angle-of-attack. In addition, the measurements produced estimates of the rigid-body blade pitch, flap, and lag angles.

rotor blade displacement↗

NASA NTRS2021

Validation of IASI Satellite Ammonia Observations at the Pixel Scale Using In Situ Vertical Profiles

Satellite ammonia (NH 3 ) observations provide unprecedented insights into NH3 emissions, spatiotemporal variabilities and trends, but validation with in situ measurements remains lacking. Here, total columns from the Infrared Atmospheric Sounding Interferometer (IASI) were intercompared to boundary layer NH 3 profiles derived from aircraft- and surface-based measurements primarily in Colorado, USA, in the summer of 2014. IASI-NH 3 version 3 near real-time data set compared well to in situ derived columns (windows ±15 km around centroid, ±1 h around overpass time) with a correlation of 0.58, a slope of 0.78 ± 0.14 and an intercept of 2.1 × 10 15 ±1.5 × 10 15 molecules cm −2 . Agreement degrades at larger spatiotemporal windows, consistent with the short atmospheric lifetime of NH 3 . We also examined IASI version 3R data, which relies on temperature retrievals from the ERA Reanalysis, and a third product generated using aircraft-measured temperature profiles. The overall agreement improves slightly for both cases, and neither is biased within their combined measurement errors. Thus, spatiotemporal averaging of IASI over large windows can be used to reduce retrieval noise. Nonetheless, sampling artifacts of airborne NH 3 instruments result in significant uncertainties of the in situ-derived columns. For example, large validation differences exist between ascent and descent profiles, and the assumptions of the free tropospheric NH 3 profiles used above the aircraft ceiling significantly impact the validation. Because short-lived species like NH 3 largely reside within the boundary layer with complex vertical structures, more comprehensive validation is needed across a wide range of environments. More accurate and widespread in situ NH 3 data sets are therefore required for improved validations of satellite products.

Ammonia↗

NASA NTRS2020

1- by 1-Foot Supersonic Wind Tunnel Sidewall Flowfield Compression Plate Study

A method for locally increasing the dynamic pressure in the test section of a supersonic wind tunnel is investigated from a theoretical, computational and experimental perspective, culminating in a reduced-scale, proof-of-concept test at Mach 3.44. Symmetric, variable-angle compression plates along the test section sidewalls generate oblique shockwaves to elevate the test section dynamic pressure (q). The experimental effort utilizes boundary-layer pitot rakes to sample the incoming flow near the floor, pitot and static rakes to measure the high-q core flow and numerous surface taps to map out the static pressure distribution along the floor, ceiling and sidewalls. A reliable and robust starting procedure is developed to establish supersonic flow in the test section and behind the compression plates. A characterization of the incoming floor boundary layer reveals previously documented features (such as a bulge toward the centerline) and an unexpected (and as-yet-unexplained) thickening of the boundary layer with increasing Reynolds number around total pressures of 40–60 psia. The high-q core flow is characterized for eight compression plate angles (4.0–8.0°), eight tunnel total pressures (20–139 psia) and three axial stations (24.95–27.35 inches). The peak dynamic pressure rise exceeds expectations. At conditions relevant to full-scale testing (6.0° and 139 psia), the dynamic pressure increases by a factor of 1.73, the Mach number becomes 2.85 and the total pressure drops imperceptibly. Basic distortion metrics show transverse non-uniformity of less than 5 %. Limited spatial resolution in the static rake measurements warrants additional, higher fidelity testing or computational analysis. Surface-mounted, fast-response instrumentation suggests minimal unsteadiness in the flow, less than a few percent of the mean (comparable to the sensor accuracy).

Wind tunnel testing, test techniques/methods, dyna↗

NASA NTRSDate not supplied

Blade Displacement Measurements of a Rotor in Forward Flight in the Langley 14-by 22-Foot Wind Tunnel

Stereo photogrammetry was used to measure the elastic bending and twist of a three-bladed rotor in forward flight in the NASA Langley 14-by 22-Ft Subsonic Wind Tunnel. The rotor was imaged from below by two cameras mounted at windows of a large turntable in the floor of the test section. At each test condition, the turntable and cameras were rotated and the cameras were triggered to allow measurements over a range of blade azimuth angles. Retro-reflective targets were applied along the leading and trailing edges of the lower surface of each blade. Image blur due to blade motion was minimized by illuminating the targets with high-intensity LED flash lamps mounted next to each camera. The cameras were re-calibrated at each angular position of the turntable using the positions in each image of retro-reflective targets on the ceiling of the test section whose spatial coordinates had been very accurately measured. The blade displacement measurements yielded elastic bending and twist along each blade as functions of blade azimuth, thrust coefficient, and advance ratio, all at a constant rotor shaft angle. In addition, the measurements produced estimates of the rigid-body blade pitch, flap, and lag angles.

rotor↗

NASA NTRS2018

Modeling Key Predictors of Airport Runway Configurations Using Learning Algorithms

Advanced traffic flow management automation will need accurate predictions of airport runway configurations. Terminal area weather and traffic demand are generally considered to be the most significant factors in predicting runway configuration. Weather information is forecasted across multiple features, including wind direction, wind speed, gusts, cloud ceilings, visibility, temperature, and precipitation, among many others. We use machine learning techniques on historical weather and runway data to determine weather features that correlate well with runway configurations. We analyze the predictive capability of weather features using different learning models trained on data from four major U.S. airports: Atlanta (ATL), Washington – Dulles (IAD), New York – Kennedy (JFK), and San Francisco (SFO). Wind direction alone is strongly correlated with runway configurations above all other examined factors, as expected. This correlation is the most significant component of the ~80% prediction accuracy in selecting between the two most frequently used runway configurations. However, individual airports show variations on how well the runway configuration decisions correlate with wind direction. While wind direction was identified as the most significant indicator of configuration decisions in ATL, IAD, and JFK, it did not emerge as such at SFO. Traffic demand was not found to be a strong factor in predicting runway configurations at any of the airports analyzed. In rare instances, when high demand cannot be accommodated within the current configuration, temporary changes are likely to be attributable to demand. However, these occurrences are so limited in number that their overall effect is not sufficient to consider traffic demand as a major indicator of runway configuration at the airports analyzed.

Bilimoria, Karl D.↗

NASA NTRSDate not supplied

Internal Architecture of the Common Habitat

The core stage liquid oxygen tank of the Space Launch System can be manufactured as a habitat instead of as a propellant tank, with a common design such that it is equally suitable for use in 0g, 1/6g, 3/8g, 1g, or variable artificial gravity. This Common Habitat can be the central element of a human spaceflight architecture that encompasses the Moon, Mars, and other destinations within the inner solar system. The Common Habitat is specifically used within an alternative architecture study (that is not part of the current NASA baseline) as the core habitation element within a Lunar Basecamp, Mars Basecamp, and the Deep Space Exploration Vehicle. The Common Habitat internal architecture applies a design philosophy to separate crew functions according to deck. The lower deck is reserved for private functions. It includes eight private crew quarters and four waste and hygiene clusters – each with a private waste management compartment, private full body hygiene compartment, and foyer/clothes changing area. The mid deck is primarily allocated to mission-related working functions. It includes an exercise facility, fabrication / maintenance / repair facility, physical science laboratory (physics, geology, and astronomy), and life science laboratory (biology and human research). The mid deck also has four external hatches, clocked one every 90 degrees, centered on the vehicle vertical centerline. Each hatch has a 60-inch tall by 40-inch wide opening with the mid deck floor 16 inches below the bottom lip of the hatch opening. The upper deck is primarily allocated to group or social functions. It includes a large galley, wardroom with projector and display screen, plant growth chambers, bulk stowage, command and control station, medical facility, hygiene compartment, and vehicle subsystems. A Vertical Translation System, based in part on results from a GrabCAD public challenge, enables crew and equipment traverse between habitat decks in 0g, 1/6g, 3/8g, and 1g. A 40-inch by 60-inch vertical translation corridor ensures that anything that can fit through the external hatches can also be moved from deck to deck. The system includes three component systems: a deploying floor to safety barrier system that folds down to become a traversable floor when vertical translation is not in use and deploys to become a safety barrier during vertical translations; a folding ladder that stows in the ceiling when not in use; and an elevator platform that can attach to the ladder for cargo or incapacitated crew member transport. This system is complemented by a Horizontal Translation System, also based in part on the same GrabCAD challenge, provides for horizontal crew translation in microgravity but requires no vehicle reconfiguration for fractional or whole gravity operation. It relies primarily on gecko-inspired grippers, originally developed at Stanford and NASA JPL, attached to footwear to enable crew to emulate walking. Some handrails are also strategically placed throughout the Common Habitat to aid translation in both gravity and microgravity environments.

Common Habitat↗

NASA NTRS2022

Assessment of the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel Supersonic Test Section for Sonic Boom and Supersonic Testing

The NASA Glenn Research Center performed a Sonic Boom Exploratory test with Boeing as an industry partner. The test was performed in the transonic test section of the 8-by 6-foot supersonic wind tunnel (8x6 SWT) in 2012. Since that test, research has been performed to validate the suitability of the supersonic test section, located upstream of the 8x6 transonic test section. This little used upstream supersonic test section can be better for certain objectives in supersonic testing and sonic boom validation. Background pressure profiles were collected in a Mach number stability test (in 2014) with a sonic boom pressure measurement rail installed in the ceiling of the smooth supersonic test section. Results from this test demonstrated favorable background pressure profiles for future sonic boom validation tests. A preliminary supersonic test section calibration was performed in 2016, followed by a detailed calibration in 2020; which also demonstrated the stable and uniform flow characteristics in the supersonic test section. The calibration test utilized the transonic array and the cone-cylinder calibration model in the smooth supersonic test section of the 8x6 SWT. Completion of the calibration prepared the facility for sonic boom assessments of the Low Boom Flight Demonstrator.

Sonic Boom↗

NASA NTRSDate not supplied

Runway Configuration Management with Offline Reinforcement Learning

Runway configuration management (RCM) is a challenging task, and it affects the efficiency of the National Airspace System (NAS) and airport surface operations significantly. Each airport, depending on the geometry, capacity, local climate patterns, etc. has multiple configurations for the runway usage for arriving and departing flights. Many factors such as the incoming/outgoing traffic load, wind direction and speed, convective weather, cloud ceiling and other environmental factors might affect the choice of a runway configuration at any point in time. However, other factors such as safety measures and regulations, noise abatement, capacity of each configuration, and preference of the air traffic controllers (ATCs) can also play a significant role in selecting the configuration. A sub-optimal selection of the runway configuration, or delay in making configuration changes might result in significant increase in taxi times for aircraft on the surface of the airport, fuel and energy use of the aircraft, and maintenance costs. It can also lead to safety concerns, such as an aircraft performing one or more go-arounds before being able to land. All these factors make RCM an extremely important and challenging decision-making process for the ATCs. The current state of practice sets the runway configuration by the ATCs based on relevant information available at the time including weather, traffic, noise abatement, safety bounds, etc. This makes the decision-making process subjective based on the accuracy of the available information and the bias in human decision making. Unfortunately, this approach yields poor results (e.g., significant delays) if the predicted outcomes are uncertain and their relative impact is not well understood. This is especially evident when the uncertainty increases the size of possible predicted outcomes (combinatorial explosion in possible scenarios) that cannot be handled by human reasoning. On the other hand, an automated approach based on machine intelligence can make use of historical data and search through all (or significant amount of) possible scenarios under uncertainty and make well-informed decisions.

Milad Memarzadeh↗

NASA NTRSDate not supplied

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper presents and discusses the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of the test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model angle of incidence. Laser Doppler velocimetry (LDV) measurements were made at three model angles of incidence (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data were obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data on the test article and Reynolds-averaged Navier-Stokes CFD results are presented and discussed.

Juncture Flow↗

NASA NTRS2023

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper will present and discuss the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of that test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model pitch angle. Laser Doppler velocimetry (LDV) measurements were made at three model pitch angles (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data was obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data and Reynolds averaged Navier-Stokes CFD results will also be presented and discussed.

Juncture Flow↗

NASA NTRSDate not supplied

Lessons Learned From Designing and Building Low Organic Outgassing Cleanrooms for NASA OSIRIS-REx and Hayabusa2 Curation Facility

NASA Johnson Space Center Astromaterials Acquisition and Curation Office completed construction and commissioning of the OSIRIS-REx and Hayabusa2 cleanroom laboratory suites along with new precision cleaning and advanced curation laboratories consisting of ISO Class 5, 6, and 7 cleanrooms. The new curation facilities are designed for initial receiving, basic characterization, curation processing, and preliminary examination of carbonaceous asteroidal material. The facilities are also designed to enable long-term pristine sample storage to preserve the scientific integrity of each sample to enable future research by the international science community. The scientific study of organics is critical for both missions. The OSIRIS-REx mission instituted a stringent contamination control plan for low organics. Given these mission requirements and long-term storage preservation requirements, cleanroom construction materials were carefully selected to not hinder the scientific search for amino acids and the study of organics in the samples. Cleanroom candidate construction materials were researched for the HVAC system, floors, walls, ceiling, and plenum areas. The team chose several candidate materials that were further tested for their specific outgassing characteristics using ASTM E-595 per ASTM E-2312. The information gained provided the foundation for the final design selection of the cleanroom materials. While material selection is important before and during the facility design phase, the construction phase is where these choices are implemented by the prime construction contractor through final product and material submittals. These submittals are the last check point and changes are often submitted due to engineering/construction conflicts, especially when renovating an existing building. Any new material or equipment/product change must be identified and scrutinized for low particulate shedding and outgassing properties. In some cases, quick testing of new material is required, and compromises must be made in real-time. Careful selection and implementation of cleanroom materials significantly reduced organic and inorganic contamination beyond normal cleanroom baselines.

astromaterials curation↗

NASA NTRSDate not supplied

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper will present and discuss the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of that test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model pitch angle. Laser Doppler velocimetry (LDV) measurements were made at three model pitch angles (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data was obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data and Reynolds averaged Navier-Stokes CFD results will also be presented and discussed.

Juncture Flow↗

NASA NTRSDate not supplied

Lessons Learned From Designing and Building Low Organic Outgassing Cleanrooms for NASA OSIRIS-REx and Hayabusa2 Curation Facility

astromaterials curation↗

NASA NTRSDate not supplied

Design and Technology Maturation of the Stratospheric Projectile Experiment of Entry Dynamics

The supersonic and transonic dynamic stability of blunt-body reentry vehicles currently poses large risks in all of NASA’s ongoing entry missions (MSR SRL, MSR EES, and Dragonfly). These projects have allocated millions of dollars to testing and modeling efforts to buy down risk by using the current state-of-the-art (SoA) facilities at NASA’s disposal. While these facilities have heritage in supplying dynamics data to reentry missions, their availability is severely limited – particularly with the high number of concur-rent projects requesting simultaneous testing– and are costly when considering the science density per dollar. None of the current SoA facility methodologies allow the test model to have the dynamics fully develop through a flight relevant free-stream profile and as such require extrapolations with resultant high uncertainties in order to relate the test dynamics to flight expectations. SPEED is a NASA Ames Center Innovation Fund (CIF) project that is developing a highly tailorable and cost-effective test methodology to better assess the dynamic stability of blunt-body reentry vehicles via a stratospheric balloon flight. This is accomplished by dropping a suite of instrumented capsules from a stratospheric balloon to gain a statistically relevant dataset of scaled reentry vehicles in mission relevant free-flight conditions. This presentation will walk through how the test methodology is being implemented specifically for the Mars Sample Return (MSR) Earth Entry System (EES) geometry in an awarded Flight Opportunities Program (FOP) test flight in early CY24. SPEED Application to MSR: SPEED consists of three main mechanical systems: the Drop Platform, the Projectile, and the test Capsule. SPEED is being developed as a set of guidelines and recommendations for how to test with the proposed Concept of Operations (Conops) since the specific design parameters will vary depending on the specific project’s reference trajectory and entry vehicle design. As such, this presentation will walk through the development time-line as shown in Fig. 2. This is meant to serve as a blueprint for further missions as desired. Mechanical and Avionics Design. The SPEED test platform designed for the MSR-EES capsule geometry with nominal entry parameters has the ability to carry 10 Capsules to altitude instrumented with: 1. 3-Axis Accelerometer 2. IMU 3. Gyroscope 4. Magnetometer 5. Pressure Transducer cruciform 6. Uplook and Horizon Cameras To package the avionics/instrumentation suite, the capsule is approximately 1’ in diameter with the Outer Mold Line (OML) centroid-scaled from the full EES design. The internal volume is gutted and custom-shaped to fit the desired instrumentation suite, as well as to allow for the positioning of ballast mass such that the Center of Gravity is analogous to the flight vehicle. All structural components in the Capsule and Projectile are 3D printed, which significantly reduces the cost of each flight unit to around $1500 including all instrumentation, avionics, and structural components. Flight Conops. The test Capsule is accelerated to the desired altitude and Mach number while stowed in the Projectile, a missile-like vehicle consisting of steel ballast in the nose, a low-drag OML, and an Ejection Mechanism to reliably release the Capsule into the free-flow supersonic conditions. For the MSR-EES design, the capsule employs ~3kg of ballast mass at the nose to accelerate the 1.25kg test Capsule to ~Mach 1.7 at 23km altitude. This requires an initial release altitude of 40km, the quoted limit of a 80kg payload by the FOP-contracted balloon provider. Once the Ejection Mechanism avionics detect the proper conditions, the spring-loaded Ejection Mechanism will release and – guided by the sabot – expose the test Capsule to the desired test conditions for ~5 seconds of free-flight in the supersonic/transonic regimes. Dynamics in the subsonic regime will also be captured with the instrumentation suite with post-flight recovery operations aimed at recovering the high-G-load capable SD cards after the planned hard impact landings. Testing and Development: In the few months the SPEED project has worked the development of MSR-EES flight test, the team has performed lab and drone based testing which this presentation will overview. After the first design phase, the team fabricated Engineering Demonstration Units (EDUs) of all subsystems to perform validation testing shown in Fig. 5. After validation was completed on the subsystem level, a drone-drop test was performed at the recreational flight ceiling of 400ft altitude to assess the SPEED systems in a flight environment. Parameters such as in-flight stability, hard impact landing performance, and avionics performance were quantified and qualified. The FY23 CIF will culminate in a helicopter drop test aboard an Air National Guard Blackhawk. This will prepare the team for the CY24 FOP stratospheric balloon flight that should provide the final verification to begin offering the test platform for mission support. Focus of Presentation: This presentation will outline the technology maturation path of the SPEED implementation to the MSR-EES capsule baseline as well as the details regarding the mechanical system, avionics and instrumentation, and flight operations. Note that a complementary presentation is being submitted for a methodology overview of the SPEED test platform, introducing the testing technique and benefits as well as the full application space of the technology.

pitch damping coefficient↗

NASA NTRSDate not supplied

Applying Machine Learning Tools for Runway Configuration Decision Support

Determining optimal runway configurations at airports, a responsibility assigned to air traffic controllers, is a challenging task. The decision-making process is intricate and involves consideration of many factors such as prevailing wind condition, convective weather, visibility, cloud ceilings, departure and arrival demand, traffic flow, equipment status, and other airport constraints. In a previous work, we developed a Runway Configuration Assistance tool using an offline reinforcement learning method called conservative Q-learning. In this paper, we evaluate and validate our Runway Configuration Assistance tool as a decision support for air traffic controllers. We validated our tool using three airports with differing levels of complexity: Charlotte Douglas International Airport, Denver International Airport, and Dallas Fort Worth International Airport. We quantified the performance of the Runway Configuration Assistance tool based on (1) agreement with historical air traffic controller decisions and (2) violation of decisions that would be obvious to subject-matter experts. Our tool showed promising results in both performance metrics for the three airports, despite the complexities in the runway configuration decision-making process. We also discuss challenges in using machine learning in general to aid air traffic management and identify deployment considerations for the Runway Configuration Assistance tool.

Runway Configuration Management↗

NASA NTRSDate not supplied

Machine Learning Tools for Runway Configuration Decision Support

Runway Configuration Management↗

NASA NTRS2024

ECI Final Report: Closing of Critical Technology Gaps for Rotating Detonation Rocket Engines

NASA has successfully fired several configurations of a dual regenerative 10K lbf class Rotating Detonation Rocket Engines using cryogenic liquid methane, liquid hydrogen, and RP-1 as fuel and liquid or gaseous oxygen as the oxidizer. In addition, 3 hot fire test phases were conducted to evaluate parametrically varying combustor geometries using a subscale 500-1K lbf class thruster RDRE. Several studies were successfully completed to evaluate the impacts of contraction ratio, annulus L’, heat transfer through calorimetry, and varying injector designs. Laser powder bed fusion GRCop-42, GRX-810, and C-103 alloys were used for the outer bodies, inner bodies, nozzles, and injector variations. The development of RDRE technology is only now possible with additive manufacturing techniques and these specialized alloys. Critical design parameters, design rules of thumb, and operability limitations have been identified. Major lessons learned in the development of critical components have also been documented. All critical technology gaps have been demonstrated at one scale or another and closed. A major finding of this work is that the combustion efficiency of the RDRE is far superior to that of the traditional liquid rocket engine requiring an order of magnitude lower chamber volume and residence time to react. Because of this, the design trade space opens up significantly. More compact chamber geometries are now possible with potential mass savings and significant length savings possible to flight geometries. In addition, the ceiling for the technology, in terms of Isp, has not yet been reached. It is likely that an additional ~10% better Isp over equivalent state-of-the-art combustion devices may be achieved over the next decade should the technology continue to be developed. This technology has been shown to be viable for near term infusion (2-5 years) into space missions and enable cost effective space access for US industry. This final report documents all major findings from this early career initiative (ECI) project and lays out what is needed post ECI.

Thomas Teasley↗

NASA NTRS2025

Optimal Mission Profile for Low-Boom Supersonic Aircraft

An optimal mission profile, called the low-boom mission profile, is proposed for low-boom supersonic aircraft. The low-boom mission profile uses a fixed AoA for the cruise segment until reaching the cruise ceiling. The fixed AoA is determined by the cruise condition for the low-boom shaping of the aircraft. The derived CL trajectory is not necessarily the optimal trajectory for the cruise range when unrestricted supersonic flight is allowed. The undertrack sonic boom ground noise level is reduced during the cruise because of the increasing altitude from the fixed AoA trajectory and the approximate invariance of nondimensional overpressure at 3BL below a supersonic aircraft with respect to cruise altitude change when the AoA and cruise Mach are fixed. The approximate invariance of nondimensional overpressure at 3BL below a supersonic aircraft is verified using both Euler and Reynolds-averaged Navier-Stokes CFD solvers. A low-boom supersonic transport concept is used to illustrate that, compared to the fixed cruise altitude trajectory, the low-boom mission profile can increase the cruise range of a low-boom aircraft and reduce the undertrack sonic boom ground noise levels for the entire cruise segment. One significant contribution of this paper is the implication that an aircraft OML satisfying a sonic boom ground noise level constraint at SOC will satisfy the same noise level constraint over the entire cruise segment, as long as a fixed AoA trajectory is flown.

Low-boom characteristics↗