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NASA NTRS2013

Conceptual Design and Cost Estimate of a Subsonic NASA Testbed Vehicle (NTV) for Aeronautics Research

A conceptual design and cost estimate for a subsonic flight research vehicle designed to support NASA's Environmentally Responsible Aviation (ERA) project goals is presented. To investigate the technical and economic feasibility of modifying an existing aircraft, a highly modified Boeing 717 was developed for maturation of technologies supporting the three ERA project goals of reduced fuel burn, noise, and emissions. This modified 717 utilizes midfuselage mounted modern high bypass ratio engines in conjunction with engine exhaust shielding structures to provide a low noise testbed. The testbed also integrates a natural laminar flow wing section and active flow control for the vertical tail. An eight year program plan was created to incrementally modify and test the vehicle, enabling the suite of technology benefits to be isolated and quantified. Based on the conceptual design and programmatic plan for this testbed vehicle, a full cost estimate of $526M was developed, representing then-year dollars at a 50% confidence level.

Nickol, Craig L.↗

NASA NTRS2013

Multifunctional Low-Pressure Turbine for Core Noise Reduction, Improved Efficiency, and Nitrogen Oxide (NOx) Reduction

This work studied the feasibility of using Helmholtz resonator cavities embedded in low-pressure-turbine (LPT) airfoils to (1) reduce core noise by damping acoustic modes; (2) use the synthetic jets produced by the liner hole acoustic oscillations to improve engine efficiency by maintaining turbulent attached flow in the LPT at low-Reynolds-number cruise conditions; and (3) reduce engine nitrogen oxide emissions by lining the internal cavities with materials capable of catalytic conversion. Flat plates with embedded Helmholtz resonators, designed to resonate at either 3000 or at 400 Hz, were simulated using computational fluid dynamics. The simulations were conducted for two inlet Mach numbers, 0.25 and 0.5, corresponding to Reynolds numbers of 90 000 and 164 000 based on the effective chordwise distance to the resonator orifice. The results of this study are (1) the region of acoustic treatment may be large enough to have a benefit; (2) the jets may not possess sufficient strength to reduce flow separation (based on prior work by researchers in the flow control area); and (3) the additional catalytic surface area is not exposed to a high velocity, so it probably does not have any benefit.

Miller, Christopher J.↗

NASA NTRS2014

Status of Propulsion Technology Development Under the NASA In-space Propulsion Technology Program

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Hall-effect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The HEP system is composed of the High Voltage Hall Accelerator (HiVHAc) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HiVHAc are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs like: MAV propulsion and electric propulsion. And finally, one focus of the SystemsMission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

propellant tanks↗

NASA NTRS2014

Status of Propulsion Technology Development Under the NASA In-space Propulsion Technology Program

electric propulsion↗

NASA NTRS2014

Thrust Measurement of Dielectric Barrier Discharge (DBD) Plasma Actuators: New Anti-Thrust Hypothesis, Frequency Sweeps Methodology, Humidity and Enclosure Effects

We discuss thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators devices used for aerodynamic active flow control. After a review of our experience with conventional thrust measurement and significant non-repeatability of the results, we devised a suspended actuator test setup, and now present a methodology of thrust measurements with decreased uncertainty. The methodology consists of frequency scans at constant voltages. The procedure consists of increasing the frequency in a step-wise fashion from several Hz to the maximum frequency of several kHz, followed by frequency decrease back down to the start frequency of several Hz. This sequence is performed first at the highest voltage of interest, then repeated at lower voltages. The data in the descending frequency direction is more consistent and selected for reporting. Sample results show strong dependence of thrust on humidity which also affects the consistency and fluctuations of the measurements. We also observed negative values of thrust, or "anti-thrust", at low frequencies between 4 Hz and up to 64 Hz. The anti-thrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic anti-thrust curve are correlated with appearance of visible plasma discharges. We propose the anti-thrust hypothesis. It states that the measured thrust is a sum of plasma thrust and anti-thrust, and assumes that the anti-thrust exists at all frequencies and voltages. The anti-thrust depends on actuator geometry and materials and on the test installation. It enables the separation of the plasma thrust from the measured total thrust. This approach enables more meaningful comparisons between actuators at different installations and laboratories. The dependence on test installation was validated by surrounding the actuator with a grounded large-diameter metal sleeve. Strong dependence on humidity is also shown; the thrust significantly increased with decreasing humidity, e.g., 44 percent increase as relative humidity changed from 18 percent and dew point 33 degF to 50 percent and dew point of 57 degF.

Ashpis, David E.↗

NASA NTRS2014

Noise and Fuel Burn Reduction Potential of an Innovative Subsonic Transport Configuration

A study is presented for the noise and fuel burn reduction potential of an innovative double deck concept aircraft with two three-shaft direct-drive turbofan engines. The engines are mounted from the fuselage so that the engine inlet is over the main wing. It is shown that such an aircraft can achieve a cumulative Effective Perceived Noise Level (EPNL) about 28 dB below the current aircraft noise regulations of Stage 4. The combination of high bypass ratio engines and advanced wing design with laminar flow control technologies provide fuel burn reduction and low noise levels simultaneously. For example, the fuselage mounted engine position provides more than 4 EPNLdB of noise reduction by shielding the inlet radiated noise. To identify the potential effect of noise reduction technologies on this concept, parametric studies are presented to reveal the system level benefits of various emerging noise reduction concepts, for both engine and airframe noise reduction. These concepts are discussed both individually to show their respective incremental noise reduction potential and collectively to assess their aggregate effects on the total noise. Through these concepts approximately about 8 dB of additional noise reduction is possible, bringing the cumulative noise level of this aircraft to 36 EPNLdB below Stage 4, if the entire suite of noise reduction technologies would mature to practical application. In a final step, an estimate is made for this same aircraft concept but with higher bypass ratio, geared, turbofan engines. With this geared turbofan propulsion system, the noise is estimated to reach as low as 40-42 dB below Stage 4 with a fuel burn reduction of 43-47% below the 2005 best-in-class aircraft baseline. While just short of the NASA N+2 goals of 42 dB and 50% fuel burn reduction, for a 2025 in service timeframe, this assessment shows that this innovative concept warrants refined study. Furthermore, this design appears to be a viable potential future passenger aircraft, not only in meeting the regulatory requirements, but also in competing with aircraft of different advanced designs within this N+2 timeframe and goal framework.

Guo, Yueping↗

NASA NTRS2014

The Status of Spacecraft Bus and Platform Technology Development Under the NASA ISPT Program

The In-Space Propulsion Technology (ISPT) program is developing spacecraft bus and platform technologies that will enable or enhance NASA robotic science missions. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems Mission Analysis. ISPTs propulsion technologies include: 1) NASAs Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; 2) a Hall-effect electric propulsion (HEP) system for sample return and low cost missions; 3) the Advanced Xenon Flow Control System (AXFS); ultra-lightweight propellant tank technologies (ULTT); and propulsion technologies for a Mars Ascent Vehicle (MAV). The AXFS and ULTT are two component technologies being developed with nearer-term flight infusion in mind, whereas NEXT and the HEP are being developed as EP systems. ISPTs entry vehicle technologies are: 1) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GNC) models of blunt-body rigid aeroshells; and aerothermal effect models; and 2) Multi-mission technologies for Earth Entry Vehicles (MMEEV) for sample return missions. The Systems Mission Analysis area is focused on developing tools and assessing the application of propulsion, entry vehicle, and spacecraft bus technologies to a wide variety of mission concepts. Several of the ISPT technologies are related to sample return missions and other spacecraft bus technology needs like: MAV propulsion, MMEEV, and electric propulsion. These technologies, as well as Aerocapture, are more vehicle and mission-focused, and present a different set of technology development challenges. These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

electric propulsion↗

NASA NTRS2014

Status of Propulsion Technology Development Under the NASA In-Space Propulsion Technology Program

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems/Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Halleffect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The NEXT Long Duration Test (LDT) recently exceeded 50,000 hours of operation and 900 kg throughput, corresponding to 34.8 MN-s of total impulse delivered. The HEP system is composed of the High Voltage Hall Accelerator (HIVHAC) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HIVHAC are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs: MAV propulsion and electric propulsion. And finally, one focus of the Systems/Mission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

systems analysis↗

NASA NTRS2014

Dielectric Barrier Discharge (DBD) Plasma Actuators Thrust-Measurement Methodology Incorporating New Anti-Thrust Hypothesis

We discuss thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators devices used for aerodynamic active flow control. After a review of our experience with conventional thrust measurement and significant non-repeatability of the results, we devised a suspended actuator test setup, and now present a methodology of thrust measurements with decreased uncertainty. The methodology consists of frequency scans at constant voltages. The procedure consists of increasing the frequency in a step-wise fashion from several Hz to the maximum frequency of several kHz, followed by frequency decrease back down to the start frequency of several Hz. This sequence is performed first at the highest voltage of interest, then repeated at lower voltages. The data in the descending frequency direction is more consistent and selected for reporting. Sample results show strong dependence of thrust on humidity which also affects the consistency and fluctuations of the measurements. We also observed negative values of thrust or "anti-thrust", at low frequencies between 4 Hz and up to 64 Hz. The anti-thrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic anti-thrust curve are correlated with appearance of visible plasma discharges. We propose the anti-thrust hypothesis. It states that the measured thrust is a sum of plasma thrust and anti-thrust, and assumes that the anti-thrust exists at all frequencies and voltages. The anti-thrust depends on actuator geometry and materials and on the test installation. It enables the separation of the plasma thrust from the measured total thrust. This approach enables more meaningful comparisons between actuators at different installations and laboratories. The dependence on test installation was validated by surrounding the actuator with a large diameter, grounded, metal sleeve.

Ashpis, David E.↗

NASA NTRS2015

Case Studies for the Statistical Design of Experiments Applied to Powered Rotor Wind Tunnel Tests

The application of statistical Design of Experiments (DOE) to helicopter wind tunnel testing was explored during two powered rotor wind tunnel entries during the summers of 2012 and 2013. These tests were performed jointly by the U.S. Army Aviation Development Directorate Joint Research Program Office and NASA Rotary Wing Project Office, currently the Revolutionary Vertical Lift Project, at NASA Langley Research Center located in Hampton, Virginia. Both entries were conducted in the 14- by 22-Foot Subsonic Tunnel with a small portion of the overall tests devoted to developing case studies of the DOE approach as it applies to powered rotor testing. A 16-47 times reduction in the number of data points required was estimated by comparing the DOE approach to conventional testing methods. The average error for the DOE surface response model for the OH-58F test was 0.95 percent and 4.06 percent for drag and download, respectively. The DOE surface response model of the Active Flow Control test captured the drag within 4.1 percent of measured data. The operational differences between the two testing approaches are identified, but did not prevent the safe operation of the powered rotor model throughout the DOE test matrices.

Overmeyer, Austin D.↗

NASA NTRS2016

Force Measurement Improvements to the National Transonic Facility Sidewall Model Support System

The National Transonic Facility is a transonic pressurized cryogenic facility. The development of the high Reynolds number semi-span capability has advanced over the years to include transonic active flow control and powered testing using the sidewall model support system. While this system can be used in total temperatures down to -250Â F for conventional unpowered configurations, it is limited to temperatures above -60Â F when used with powered models that require the use of the high-pressure air delivery system. Thermal instabilities and non-repeatable mechanical arrangements revealed several data quality shortfalls by the force and moment measurement system. Recent modifications to the balance cavity recirculation system have improved the temperature stability of the balance and metric model-to-balance hardware. Changes to the mechanical assembly of the high-pressure air delivery system, particularly hardware that interfaces directly with the model and balance, have improved the repeatability of the force and moment measurement system. Drag comparisons with the high-pressure air system removed will also be presented in this paper.

Goodliff, Scott L.↗

NASA NTRS2016

Dynamic Stall Suppression Using Combustion-Powered Actuation (COMPACT)

Retreating blade stall is a well-known phenomenon that limits rotorcraft speed, maneuverability, and efficiency. Airfoil dynamic stall is a simpler problem, which demonstrates many of the same flow phenomena. Combustion Powered Actuation (COMPACT) is an active flow control technology, which at the outset of this work, had been shown to mitigate static and dynamic stall at low Mach numbers. The attributes of this technology suggested strong potential for success at higher Mach numbers, but such experiments had never been conducted. The work detailed in this report documents a 3-year effort focused on assessing the effectiveness of COMPACT for dynamic stall suppression at freestream conditions up to Mach 0.5. The work done has focused on implementing COMPACT on a high-lift rotorcraft airfoil: the VR-12. This selection was made in order to ensure that any measured benefits are over and above the capabilities of state-of-the-art high-lift rotorcraft airfoils. The detailed Computational Fluid Dynamics (CFD) simulations, wind-tunnel experiments, and system-level modeling conducted have shown the following: (1) COMPACT, in its current state of development, is capable of reducing the adverse effects of deep dynamic stall at Mach numbers up to 0.4; (2) The two-dimensional (2D) CFD results trend well compared to the experiments; and (3) Implementation of the CFD results into a system-level model suggest that significant rotor-level benefits are possible.

Matalanis, Claude G.↗

NASA NTRS2017

Radiation Hardening by Software Techniques on FPGAs: Flight Experiment Evaluation and Results

We present our work on implementing Radiation Hardening by Software (RHBSW) techniques on the Xilinx Virtex5 FPGAs PowerPC 440 processors on the SpaceCube 2.0 platform. The techniques have been matured and tested through simulation modeling, fault emulation, laser fault injection and now in a flight experiment, as part of the Space Test Program- Houston 4-ISS SpaceCube Experiment 2.0 (STP-H4-ISE 2.0). This work leverages concepts such as heartbeat monitoring, control flow assertions, and checkpointing, commonly used in the High Performance Computing industry, and adapts them for use in remote sensing embedded systems. These techniques are extremely low overhead (typically <1.3%), enabling a 3.3x gain in processing performance as compared to the equivalent traditionally radiation hardened processor. The recently concluded STP-H4 flight experiment was an opportunity to upgrade the RHBSW techniques for the Virtex5 FPGA and demonstrate them on-board the ISS to achieve TRL 7. This work details the implementation of the RHBSW techniques, that were previously developed for the Virtex4-based SpaceCube 1.0 platform, on the Virtex5-based SpaceCube 2.0 flight platform. The evaluation spans the development and integration with flight software, remotely uploading the new experiment to the ISS SpaceCube 2.0 platform, and conducting the experiment continuously for 16 days before the platform was decommissioned. The experiment was conducted on two PowerPCs embedded within the Virtex5 FPGA devices and the experiment collected 19,400 checkpoints, processed 253,482 status messages, and incurred 0 faults. These results are highly encouraging and future work is looking into longer duration testing as part of the STP-H5 flight experiment.

Hybrid Flight Architectures↗

NASA NTRS2017

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.↗

NASA NTRS2015

Low-Power Operation and Plasma Characterization of a Qualification Model SPT-140 Hall Thruster for NASA Science Missions

Hall thruster systems based on commercial product lines can potentially lead to lower cost electric propulsion (EP) systems for deep space science missions. A 4.5-kW SPT-140 Hall thruster presently under qualification testing by SSL leverages the substantial heritage of the SPT-100 being flown on Russian and US commercial satellites. The Jet Propulsion Laboratory is exploring the use of commercial EP systems, including the SPT-140, for deep space science missions, and initiated a program to evaluate the SPT-140 in the areas of low power operation and thruster operating life. A qualification model SPT-140 designated QM002 was evaluated for operation and plasma properties along channel centerline, from 4.5 kW to 0.8 kW. Additional testing was performed on a development model SPT-140 designated DM4 to evaluate operation with a Moog proportional flow control valve (PFCV). The PFCV was commanded by an SSL engineering model PPU-140 Power Processing Unit (PPU). Performance measurements on QM002 at 0.8 kW discharge power were 50 mN of thrust at a total specific impulse of 1250 s, a total thruster efficiency of 0.38, and discharge current oscillations of under 3% of the mean current. Steady-state operation at 0.8 kW was demonstrated during a 27 h firing. The SPT-140 DM4 was operated in closed-loop control of the discharge current with the PFCV and PPU over discharge power levels of 0.8-4.5 kW. QM002 and DM4 test data indicate that the SPT-140 design is a viable candidate for NASA missions requiring power throttling down to low thruster input power.

Garner, Charles E.↗

NASA NTRS2017

13kW Advanced Electric Propulsion Flight System Development and Qualification

The next phase of robotic and human deep space exploration missions is enhanced by high performance, high power solar electric propulsion systems for large-scale science missions and cargo transportation. Aerojet Rocketdynes Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS includes a magnetically-shielded, long-life Hall thruster, power processing unit (PPU), xenon flow controller (XFC), and intrasystem harnesses. The Hall thruster, originally developed and demonstrated by NASAs Glenn Research Center and the Jet Propulsion Laboratory, operates at input powers up to 12.5kW while providing a specific impulse over 2600s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95 to 140V, consistent with operation beyond the orbit of Mars. The integrated system is continuously throttleable between 3 and 13.3kW. The program has completed the system requirement review; the system, thruster, PPU and XFC preliminary design reviews; development of engineering models, and initial system integration testing. This paper will present the high power AEPS capabilities, overall program and design status and the latest test results for the 13.3kW flight system development and qualification program.

Jackson, Jerry↗

NASA NTRS2018

Design of a Modular Monolithic Implicit Solver for Multi-Physics Applications

The design of a modular multi-physics high-order space-time finite-element framework is presented together with its extension to allow monolithic coupling of different physics. One of the main objectives of the framework is to perform efficient high- fidelity simulations of capsule/parachute systems. This problem requires simulating multiple physics including, but not limited to, the compressible Navier-Stokes equations, the dynamics of a moving body with mesh deformations and adaptation, the linear shell equations, non-re effective boundary conditions and wall modeling. The solver is based on high-order space-time - finite element methods. Continuous, discontinuous and C1-discontinuous Galerkin methods are implemented, allowing one to discretize various physical models. Tangent and adjoint sensitivity analysis are also targeted in order to conduct gradient-based optimization, error estimation, mesh adaptation, and flow control, adding another layer of complexity to the framework. The decisions made to tackle these challenges are presented. The discussion focuses first on the "single-physics" solver and later on its extension to the monolithic coupling of different physics. The implementation of different physics modules, relevant to the capsule/parachute system, are also presented. Finally, examples of coupled computations are presented, paving the way to the simulation of the full capsule/parachute system.

Carton De Wiart, Corentin↗

NASA NTRS2017

Lattice Boltzmann for Airframe Noise Predictions

Increase predictive use of High-Fidelity Computational Aero- Acoustics (CAA) capabilities for NASA's next generation aviation concepts. CFD has been utilized substantially in analysis and design for steady-state problems (RANS). Computational resources are extremely challenged for high-fidelity unsteady problems (e.g. unsteady loads, buffet boundary, jet and installation noise, fan noise, active flow control, airframe noise, etc) ü Need novel techniques for reducing the computational resources consumed by current high-fidelity CAA Need routine acoustic analysis of aircraft components at full-scale Reynolds number from first principles Need an order of magnitude reduction in wall time to solution!

Barad, Michael↗