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At least 181 records · Page 10

Wind Tunnel Investigation of Passive Vortex Control and Vortex-Tail Interactions on a Slender Wing at Subsonic and Transonic Speeds

A wind tunnel experiment was conducted in the NASA Langley 8-Foot Transonic Pressure Tunnel to determine the effects of passive porosity on vortex flow interactions about a slender wing configuration at subsonic and transonic speeds. Flow-through porosity was applied in several arrangements to a leading-edge extension, or LEX, mounted to a 65-degree cropped delta wing as a longitudinal instability mitigation technique. Test data were obtained with LEX on and off in the presence of a centerline vertical tail and twin, wing-mounted vertical fins to quantify the sensitivity of the aerodynamics to tail placement and orientation. A close-coupled canard was tested as an alternative to the LEX as a passive flow control device. Wing upper surface static pressure distributions and six-component forces and moments were obtained at Mach numbers of 0.50, 0.85, and 1.20, unit Reynolds number of 2.5 million, angles of attack up to approximately 30 degrees, and angles of sideslip to +/-8 degrees. The off-surface flow field was visualized in cross planes on selected configurations using a laser vapor screen flow visualization technique. Tunnel-to-tunnel data comparisons and a Reynolds number sensitivity assessment were also performed. 15.

Erickson, Gary E.↗

Compact Multi-Gas Monitor for Life Support Systems Control in Space: Evaluation Under Realistic Environmental Conditions

Advanced space life support systems require lightweight, low-power, durable sensors for monitoring critical gas components. A luminescence-based optical flow-through cell to monitor carbon dioxide, oxygen, and humidity has been developed and was demonstrated using bench top instrumentation under environmental conditions relevant to portable life support systems, including initially pure oxygen atmosphere, pressure range from 3.5 to 14.7 psi, temperature range from 50 F to 150 F, and humidity from dry to 100% RH and under liquid water saturation. This paper presents the first compact readout unit for these optical sensors, designed for the volume, power, and weight restrictions of a spacesuit portable Life support system and the analytical characterization of the optical sensors interrogated by the novel optoelectronic system. Trace gas contaminants in a space suit, originating from hardware and material off-gassing and crew member metabolism, are from many chemical families. The result is a gas mix much more complex than the pure oxygen fed into the spacesuit, which may interfere with gas sensor readings. The paper also presents an evaluation of optical sensor performance when exposed to the most significant trace gases reported to be found in spacesuits. The studies were conducted with the spacecraft maximum allowable concentrations for those trace gases and the calculated 8-hr. concentrations resulting from having no trace contaminant control system in the ventilation loop. Finally, a profile of temperature, pressure, humidity, and gas composition for a typical EVA mission has been defined, and the performance of sensors operated repeatedly under simulated EVA mission conditions has been studied.

Delgado, Jesus↗

Performance of a Supersonic Over-Wing Inlet with Application to a Low-Sonic-Boom Aircraft

Development of commercial supersonic aircraft has been hindered by many related factors including fuel-efficiency, economics, and sonic-boom signatures that have prevented over-land flight. Materials, propulsion, and flight control technologies have developed to the point where, if over-land flight were made possible, a commercial supersonic transport could be economically viable. Computational fluid dynamics, and modern optimization techniques enable designers to reduce the boom signature of candidate aircraft configurations to acceptable levels. However, propulsion systems must be carefully integrated with these low-boom configurations in order that the signatures remain acceptable. One technique to minimize the downward propagation of waves is to mount the propulsion systems above the wing, such that the wing provides shielding from shock waves generated by the inlet and nacelle. This topmounted approach introduces a number of issues with inlet design and performance especially with the highly-swept wing configurations common to low-boom designs. A 1.79%-scale aircraft model was built and tested at the NASA Glenn Research Center's 8-by 6-Foot Supersonic Wind Tunnel (8x6 SWT) to validate the configuration's sonic boom signature. In order to evaluate performance of the top-mounted inlets, the starboard flow-through nacelle on the aerodynamic model was replaced by a 2.3%-scale operational inlet model. This integrated configuration was tested at the 8x6 SWT from Mach 0.25 to 1.8 over a wide range of angles-of-attack and yaw. The inlet was also tested in an isolated configuration over a smaller range of angles-of-attack and yaw. A number of boundary-layer bleed configurations were investigated and found to provide a substantial positive impact on pressure recovery and distortion. Installed inlet performance in terms of mass capture, pressure recovery, and distortion over the Mach number range at the design angle-of-attack of 4-degrees is presented herein and compared to that at 0- degrees, as well as the isolated inlet configuration to highlight installation effects. Performance of the installed inlet fell below that of the isolated inlet at Mach numbers of 1.4 and greater. The installed inlet demonstrated adequate operability over the expected range of angles-of-attack and yaw, but did exhibit definite angle-ofattack and yaw limits at supersonic conditions. At each supersonic flight Mach number, performance parameters near zero yaw angle were relatively insensitive to yaw, but in general the yaw angle yielding best performance was non-zero and varied with angle-of-attack. Performance of the installed inlet is also presented as functions of angle-of-attack and yaw to highlight these effects. Distortion at the aerodynamic interface plane ranged between 10 and 25% at the inlet critical points over the range of flight Mach numbers tested and did not decrease significantly for the isolated inlet. Although these distortion levels would be considered high for operation with a turbine engine, the over-wing installation is likely not as significant a contributor as the low test Reynolds number. This is demonstrated by comparing CFD analysis of the isolated inlet at test scale with that at intermediate and full scales.

Performance↗

Summary and Statistical Analysis of the First AIAA Sonic Boom Prediction Workshop

A summary is provided for the First AIAA Sonic Boom Workshop held 11 January 2014 in conjunction with AIAA SciTech 2014. Near-field pressure signatures extracted from computational fluid dynamics solutions are gathered from nineteen participants representing three countries for the two required cases, an axisymmetric body and simple delta wing body. Structured multiblock, unstructured mixed-element, unstructured tetrahedral, overset, and Cartesian cut-cell methods are used by the participants. Participants provided signatures computed on participant generated and solution adapted grids. Signatures are also provided for a series of uniformly refined workshop provided grids. These submissions are propagated to the ground and loudness measures are computed. This allows the grid convergence of a loudness measure and a validation metric (dfference norm between computed and wind tunnel measured near-field signatures) to be studied for the first time. Statistical analysis is also presented for these measures. An optional configuration includes fuselage, wing, tail, flow-through nacelles, and blade sting. This full configuration exhibits more variation in eleven submissions than the sixty submissions provided for each required case. Recommendations are provided for potential improvements to the analysis methods and a possible subsequent workshop.

Park, Michael A.↗

Compact Multi-Gas Monitor for Life Support Systems Control in Space: Evaluation Under Realistic Environmental Conditions

Advanced space life support systems require lightweight, low-power, durable sensors for monitoring critical gas components. A luminescence-based optical flow-through cell to monitor carbon dioxide, oxygen, and humidity has been developed and was demonstrated using bench-top instrumentation under environmental conditions relevant to portable life support systems, including initially pure oxygen atmosphere, temperature range from 50 F to 150 F, and humidity from dry to 100% RH and under conditions of water condensation. This paper presents the most recent progress in the development of this sensor technology. Trace gas contaminants in a space suit, originating from hardware and material off-gassing and crew member metabolism, are from many chemical families. The result is a gas mix much more complex than the pure oxygen fed into the space suit, and this complexity may interfere with gas sensor readings. This paper presents an evaluation of optical sensor performance when exposed to the most significant trace gases reported to be found in space suits. A study of the calibration stability of the sensors is also presented. For that purpose, a profile of temperature, pressure, humidity, and gas composition for the duration of an EVA has been defined, and the performance of sensors operated repeatedly under those conditions has been studied. Finally, this paper presents the first compact readout unit for these optical sensors, designed for the volume, power, and weight restrictions of a PLSS.

Alonso, Jesus Delgado↗

Under-Track CFD-Based Shape Optimization for a Low-Boom Demonstrator Concept

The detailed outer mold line shaping of a Mach 1.6, demonstrator-sized low-boom concept is presented. Cruise trim is incorporated a priori as part of the shaping objective, using an equivalent-area-based approach. Design work is performed using a gradient-driven optimization framework that incorporates a three-dimensional, nonlinear flow solver, a parametric geometry modeler, and sensitivities derived using the adjoint method. The shaping effort is focused on reducing the under-track sonic boom level using an inverse design approach, while simultaneously satisfying the trim requirement. Conceptual-level geometric constraints are incorporated in the optimization process, including the internal layout of fuel tanks, landing gear, engine, and crew station. Details of the model parameterization and design process are documented for both flow-through and powered states, and the performance of these optimized vehicles presented in terms of inviscid L/D, trim state, pressures in the near-field and at the ground, and predicted sonic boom loudness.

Wintzer, Mathias↗

Permeability of Two Parachute Fabrics - Measurements, Modeling, and Application

Two parachute fabrics, described by Parachute Industry Specifications PIA-C-7020D Type I and PIA-C-44378D Type I, were tested to obtain their permeabilities in air (i.e., flow-through volume of air per area per time) over the range of differential pressures from 0.146 psf (7 Pa) to 25 psf (1197 Pa). Both fabrics met their specification permeabilities at the standard differential pressure of 0.5 inch of water (2.60 psf, 124 Pa). The permeability results were transformed into an effective porosity for use in calculations related to parachutes. Models were created that related the effective porosity to the unit Reynolds number for each of the fabrics. As an application example, these models were used to calculate the total porosities for two geometrically-equivalent subscale Disk-Gap-Band (DGB) parachutes fabricated from each of the two fabrics, and tested at the same operating conditions in a wind tunnel. Using the calculated total porosities and the results of the wind tunnel tests, the drag coefficient of a geometrically-equivalent full-scale DGB operating on Mars was estimated.

Cruz, Juan R.↗

Wind Tunnel Investigation of Passive Porosity Applied to the Leading-Edge Extension and Leading-Edge Flaps on a Slender Wing at Subsonic Speed

A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.

Erickson, Gary E.↗

Permeability of Two Parachute Fabrics: Measurements, Modeling, and Application

Two parachute fabrics, described by Parachute Industry Specifications PIA-C-7020D Type I and PIA-C-44378D Type I, were tested to obtain their permeabilities in air (i.e., flow-through volume of air per area per time) over the range of differential pressures from 0.146 psf (7 Pa) to 25 psf (1197 Pa). Both fabrics met their specification permeabilities at the standard differential pressure of 0.5 inch of water (2.60 psf, 124 Pa). The permeability results were transformed into an effective porosity for use in calculations related to parachutes. Models were created that related the effective porosity to the unit Reynolds number for each of the fabrics. As an application example, these models were used to calculate the total porosities for two geometrically-equivalent subscale Disk-Gap-Band (DGB) parachutes fabricated from each of the two fabrics, and tested at the same operating conditions in a wind tunnel. Using the calculated total porosities and the results of the wind tunnel tests, the drag coefficient of a geometrically-equivalent full-scale DGB operating on Mars was estimated.

Cruz, Juan R.↗

Permeability of a New Parachute Fabric – Measurements, Modeling, and Application

In response to a NASA request, Heathcoat Fabrics Limited has woven a new parachute fabric (Custom Design 1 G-60315-1800-Q01). This fabric was tested to obtain its permeability in air (i.e., flow-through volume of air per area per time) over a range of differential pressures from 0.146 to 25 psf (7 to 1197 Pa). The fabric met its specification permeability of 60 to 100 ft3/ft2/min (30.5 to 50.8 cm3/cm2/s) at the U.S. standard differential pressure of 0.5 inch of water (2.60 psf, 124 Pa). The permeability results were transformed into an effective porosity model for use in calculations related to the total porosity of parachutes. The tested fabric is being considered for use in parachutes for future missions to Mars. Calculations of drag coefficient were performed for two geometrically identical parachutes using either the new fabric or fabric woven to Parachute Industry Specification PIA-C-7020D Type I (Mars Science Laboratory Disk-Gap-Band parachute operating on Mars at a Mach number of 0.41). These calculations indicate essentially no difference in the drag coefficient between the two parachutes.

Cruz, Juan R.↗

Development of a Non-Iterative Balance Load Prediction Algorithm for the NASA Ames Unitary Plan Wind Tunnel

A non-iterative load prediction algorithm for strain-gage balances was developed for the NASA Ames Unitary Plan Wind Tunnels that computes balance loads from the electrical outputs of the balance bridges and a set of state variables. A state variable could be, for example, a balance temperature difference or the bellows pressure of a flow-through balance. The algorithm directly uses regression models of the balance loads for the load prediction that were obtained by applying global regression analysis to balance calibration data. This choice greatly simplifies both implementation and use of the load prediction process for complex balance configurations as no load iteration needs to be performed. The regression model of a balance load is constructed by using terms from a total of nine term groups. Four term groups are derived from a Taylor Series expansion of the relationship between the load, gage outputs, and state variables. The remaining five term groups are defined by using absolute values of the gage outputs and state variables. Terms from these groups should only be included in the regression model if calibration data from a balance with known bi-directional outputs is analyzed. It is illustrated in detail how global regression analysis may be applied to obtain the coefficients of the chosen regression model of a load component assuming that no linear or massive near-linear dependencies between the regression model terms exist. Data from the machine calibration of a six-component force balance is used to illustrate both application and accuracy of the non-iterative load prediction process.

Ulbrich, Norbert M.↗

USM3D Simulations for Second Sonic Boom Workshop

The NASA Tetrahedral Unstructured Software System with the USM3D flow solver was used to compute the test cases for the Second AIAA Sonic Boom Prediction Workshop. The intent of this article is to document the USM3D results for SBPW2 test cases. The test cases included an axisymmetric equivalent area body, a JAXA wing body, a NASA low boom supersonic configuration modeled with flow-through nacelles and engine boundary conditions. All simulations were conducted for a free stream Mach number of 1.6, zero degrees angle of attack, and a Reynolds number of 5.7 million per meter. Simulations were conducted on the tetrahedral grids provided by the workshop committee, as well as a family of grids generated by an in-house approach for sonic boom analyses known as BoomGrid using current best practices. The near-field pressure signatures were extracted and propagated to the ground with the atmospheric propagation code, sBOOM. The USM3D near-field pressure signatures, corresponding sBOOM ground signatures, and loudness levels on the ground are compared with mean values from other workshop participants.

Elmiligui, Alaa↗

Testing of Mars Retropropulsion Concepts in the Langley Unitary Plan Wind Tunnel

Future Mars human lander missions will require using retrorockets beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes, the freestream flow, and the vehicle that alter the aerodynamic behavior. Historical retropropulsion ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. A new retropropulsion wind tunnel test was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to address some of the deficiencies of previous tests and to provide a basis for computational model comparisons on relevant geometries. The test included improvements in model design and measured data. Two sub-scale model geometries derived from full-scale vehicle concepts were tested: a blunt low lift-to-drag vehicle and a more slender geometry. The first model includes variations in nozzle parameters that were estimated to have first-order effects on the measurements: location, pointing direction, size, and exit-to-throat area ratio. Both models were instrumented with discrete steady and unsteady pressure sensors on the aerodynamic and interior flow path surfaces. The heats shields of both models also were covered with pressure sensitive paint that effectively measures a steady pressure distribution in an otherwise unsteady flow field. This pressure distribution was integrated over the surface geometry in order to estimate the steady aerodynamic force and moment coefficient contribution from the painted surfaces; this approach was used as a replacement for using a flow-through balance. This paper will cover the test objectives, model design, test matrix, and available test results for each model configuration at different freestream Mach numbers, model attitude, and non-dimensional thrust magnitude. A subsequent publication will cover uncertainty quantification of the test data and comparisons to computational fluid dynamics model results.

Powered Descent↗

Update on Feasibility of UV LEDs in a Spacecraft Wastewater Application

As the National Aeronautics and Space Administration (NASA) expands its scope and begins to venture into long-duration, manned space flights, the function and maintenance of spacecraft water systems becomes increasingly critical and difficult. New mission requirements limit opportunities for resupply and demand extended periods of uncrewed operations. Based on lessons learned from the International Space Station (ISS), one particular challenge of future spacecraft water systems will be maintaining adequate microbial control, especially in water subsystems and component-level elements, where effective long-duration biocontrol strategies do not currently exist. To ensure the reliability and redundancy in these systems, innovative technologies will be needed to ensure mission success. After proving feasibility of commercial-off-the-shelf (COTS) devices in flow-through applications in 2018, work focused on utilizing ultraviolet (UV) light emitting diodes (LEDs) in a wastewater tank application. This paper gives an update on the tank testing that has been performed to date. Using this technology may reduce the need for consumable resupply, such as filters or biocides, and minimize crew time to make the repairs on exhausted and/or compromised systems. The ultimate rationale behind developing a UV disinfection system is to increase the stability of water systems as requirements for sterility and microbial control become more stringent for deep space missions. The resulting data from this study will be used to narrow down possible technology demonstrations for selected ISS locations to assess the use of UV technology on future exploration-class spacecraft systems.

Susan N. Gilbert↗

Effect of Reactant Purity on Proton Exchange Membrane Fuel Cell Performance

NASA has set a goal to return to the Moon and to establish a sustained lunar presence. Many applicable lander and upperstage vehicle concepts utilize cryogenic H2 and O2 propellants. This propellant selection enables appealing potential mission concepts wherein electricity is generated by operating a fuel cell on residual H2 and O2 propellants. This concept depends on the capability of the fuel cell to utilize dry, propellant-grade reactants with concentrations of up to 30 percent He present. This study consists of the evaluation of a 12-cell non-flow-through proton exchange membrane fuel cell stack of 50 sq. cm active area with passive water removal. This stack was supplied with three levels of reactant purity: >99.999, 99.1, and 70 mol%, with the remainder made up of He. The reactant humidity and flow-through rate were also assessed as performance factors, and the fuel stack repeatedly supported a load profile with current densities up to 500 mA/sq. cm (25 A). The fuel cell performed consistently over the course of testing, with cell voltages decreasing approximately 50 mV at the maximum current density when supplied with reactants with 30 percent He present.

electrochemistry↗

Status of Mars Retropropulsion Testing in the Langley Unitary Plan Wind Tunnel

Future Mars human landings will be enabled by a powered descent phase starting at supersonic conditions, something which has never been done before on a Mars mission. Significant aerosciences challenges exist due to jet interactions between the retrorocket engine plumes, freestream flow, and vehicle that will affect the aerodynamic behavior during powered descent. Historically, wind tunnel tests have been used to study the interactions with inert gas exhaust simulants in place of rocket engines. On the computational side, flowfield simulations have been completed at full-scale conditions, but the available ground and flight data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles, due to insufficient data, dissimilar vehicle geometries, and disparate operating conditions. A wind tunnel test has been designed to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. The test will be conducted in the NASA Langley Unitary Plan Wind Tunnel and is designed with improvements in model design and data products over past tests. The test campaign will be run using sub-scale model geometries derived from NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry that generates higher unpowered lift. Both models have been fabricated and are ready for testing. The blunt model is equipped with the flexibility to examine the effects of nozzle pointing direction, number, location, size, and area ratio. The main measurements are heatshield aerodynamic interference forces and moments with a custom flow-through balance, discrete and distributed heatshield pressure, and high-speed flowfield visualization. This paper covers the test objectives, facility, models and instrumentation, and planned test matrix.

Mars↗

Status of Mars Retropropulsion Testing in the Langley Unitary Plan Wind Tunnel

Future Mars human landings will be enabled by a powered descent phase starting at supersonic conditions, something which has never been done before on a Mars mission. Significant aerosciences challenges exist due to jet interactions between the retrorocket engine plumes, freestream flow, and vehicle that will affect the aerodynamic behavior during powered descent. Historically, wind tunnel tests have been used to study the interactions with inert gas exhaust simulants in place of rocket engines. On the computational side, flowfield simulations have been completed at full-scale conditions, but the available ground and flight data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles, due to insufficient data, dissimilar vehicle geometries, and disparate operating conditions. A wind tunnel test has been designed to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. The test will be conducted in the NASA Langley Unitary Plan Wind Tunnel and is designed with improvements in model design and data products over past tests. The test campaign will be run using sub-scale model geometries derived from NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry that generates higher unpowered lift. Both models have been fabricated and are ready for testing. The blunt model is equipped with the flexibility to examine the effects of nozzle pointing direction, number, location, size, and area ratio. The main measurements are heatshield aerodynamic interference forces and moments with a custom flow-through balance, discrete and distributed heatshield pressure, and high-speed flowfield visualization. This paper covers the test objectives, facility, models and instrumentation, and planned test matrix.

Supersonic Retropropulsion↗

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↗