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At least 559 records · Page 31

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet Flows with Shock Interactions

NASA and the U.S. aerospace industry are performing studies of supersonic aircraft concepts with low sonic boom pressure signatures. The computational analyses of modern aircraft designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty remains in the aft signatures due to boundary layer and nozzle exhaust jet effects. Wind tunnel testing without inlet and nozzle exhaust jet effects at lower Reynolds numbers than in-flight make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel is planned for February 2016 to address the nozzle jet effects on sonic boom. The experiment will provide pressure signatures of test articles that replicate waveforms from aircraft wings, tails, and aft fuselage (deck) components after passing through cold nozzle jet plumes. The data will provide a variety of nozzle plume and shock interactions for comparison with computational results. A large number of high-fidelity numerical simulations of a variety of shock generators were evaluated to define a reduced collection of suitable test models. The computational results of the candidate wind tunnel test models as they evolved are summarized, and pre-test computations of the final designs are provided.

Wind Tunnel↗

HIAD on ULA (HULA) Orbital Reentry Flight Experiment Concept

This paper describes a proposed orbital velocity reentry flight test of a Hypersonic Inflatable Aerodynamic Decelerator (HIAD). The flight test builds upon ground development activities that continue to advance the materials, design, and manufacturing techniques for the inflatable structure and flexible thermal protection system (F-TPS) that comprise the inflatable heat shield. While certain aspects of material and system performance can be assessed using a variety of ground testing capabilities, only orbital velocity energy on a trajectory through the gradient density of the atmosphere can impart the combined aerodynamic and aeroheating design environments in real time. To achieve this at limited cost, the HIAD would be delivered to a spin-stabilized entry trajectory as a secondary payload on the Centaur stage of a United Launch Alliance (ULA) Atlas V launch vehicle. Initial trajectory studies indicate that the combination of launch vehicle capability and achievable reentry vehicle ballistic numbers make this a strategic opportunity for technology development. This 4 to 6 meter diameter scale aeroshell flight, referred to as HIAD on ULA (HULA), would also contribute to ULA asset recovery development. ULA has proposed that a HIAD be utilized as part of the Sensible, Modular, Autonomous Return Technology (SMART) initiative to enable recovery of the Vulcan launch vehicle booster main engines [1], including a Mid-Air Recovery (MAR) to gently return these assets for reuse. Whereas HULA will attain valuable aerothermal and structural response data toward advancing HIAD technology, it may also provide a largest-to-date scaled flight test of the MAR operation, which in turn would allow the examination of a nearly pristine post-entry aeroshell. By utilizing infrared camera imaging, HULA will also attain aft-side thermal response data, enhancing understanding of the aft side aerothermal environment, an area of high uncertainty. The aeroshell inflation will utilize a heritage design compressed gas system to minimize development costs. The data will be captured to both an onboard recorder and a recorder that is jettisoned and recovered separately from the reentry vehicle to mitigate risk. This paper provides an overview, including the architecture and flight concept of operations, for the proposed HULA flight experiment.

Dinonno, J. M.↗

Successful Completion of the JWST OGSE2 Cryogenic Test at JSC Chamber-A While Managing Numerous Challenges

The James Webb Space Telescope (JWST) Optical Telescope Element (OTE) assembly is the largest optically stable infrared-optimized telescope currently being manufactured and assembled, and scheduled for launch in 2018. The JWST OTE, including the primary mirrors, secondary mirror, and the Aft Optics Subsystems (AOS) are designed to be passively cooled and operate at near 45 degrees Kelvin. Due to the size of its large sunshield in relation to existing test facilities, JWST cannot be optically or thermally tested as a complete observatory-level system at flight temperatures. As a result, the telescope portion along with its instrument complement will be tested as a single unit very late in the program, and on the program schedule critical path. To mitigate schedule risks, a set of 'pathfinder' cryogenic tests will be performed to reduce program risks by demonstrating the optical testing capabilities of the facility, characterizing telescope thermal performance, and allowing project personnel to learn valuable testing lessons off-line. This paper describes the 'pathfinder' cryogenic test program, focusing on the recently completed second test in the series called the Optical Ground Support Equipment 2 (OGSE2) test. The JWST OGSE2 was successfully completed within the allocated project schedule while faced with numerous conflicting thermal requirements during cool-down to the final cryogenic operational temperatures, and during warm-up after the cryo-stable optical tests. The challenges include developing a pre-test cool-down and warm-up profiles without a reliable method to predict the thermal behaviors in a rarified helium environment, and managing the test article hardware safety driven by the project Limits and Constraints (L&C's). Furthermore, OGSE2 test included the time critical Aft Optics Subsystem (AOS), a part of the flight Optical Telescope Element that would need to be placed back in the overall telescope assembly integrations. The OGSE2 test requirements included the strict adherence of the project contamination controls due to the presence of the contamination sensitive flight optical elements. The test operations required close coordination of numerous personnel while they being exposed and trained for the 'final' combined OTE and instrument cryo-test in 2017. This paper will also encompass the OGSE2 thermal data look-back review.

JWST OGSE2 Cryogenic Test↗

Approach to Modeling Boundary Layer Ingestion Using a Fully Coupled Propulsion-RANS Model

Airframe-propulsion integration concepts that use boundary layer ingestion have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA's Starc-ABL aircraft configuration, which offers the potential for 12% mission fuel burn reduction by using a turbo-electric propulsion system with an aft-mounted electrically driven boundary layer ingestion propulsor. This large potential for improved performance motivates a more detailed study of the boundary layer ingestion propulsor design, but to date, analyses of boundary layer ingestion have used uncoupled methods. These methods account for only aerodynamic effects on the propulsion system or propulsion system effects on the aerodynamics, but not both simultaneously. This work presents a new approach for building fully coupled propulsive-aerodynamic models of boundary layer ingestion propulsion systems. A 1D thermodynamic cycle analysis is coupled to a RANS simulation to model the Starc-ABL aft propulsor at a cruise condition and the effects variation in propulsor design on performance are examined. The results indicates that both propulsion and aerodynamic effects contribute equally toward the overall performance and that the fully coupled model yields substantially different results compared to uncoupled. The most significant finding is that boundary layer ingestion, while offering substantial fuel burn savings, introduces throttle dependent aerodynamics effects that need to be accounted for. This work represents a first step toward the multidisciplinary design optimization of boundary layer ingestion propulsion systems.

Boundary Layer Ingestion↗

The Utility of a Small Pressurized Rover with Suit Ports for Lunar Exploration: A Geologist's Perspective

Rover trade study: As summarized recently, mission simulations at Black Point Lava Flow (Arizona) that included realistic extravehicular activity (EVA) tasking, accurate traverse timelines, and an in-loop science CAPCOM (or SciCOM) showed that a small pressurized rover (SPR) was a better mobility asset than an unpressurized rover (UPR). Traverses within the SPR were easier on crew than spending an entire day in a spacesuit, enhancing crew productivity at each station. The SPR, named Lunar Electric Rover (LER), and sometimes called the Space Exploration Vehicle (SEV), could also provide shelter during a suit malfunction, radiation event, or medical emergency that might occur on the Moon. Intravehicular activity (IVA) capabilities: From within the vehicle, crew could describe and photo-document distant features during drives between stations, as well as in the near-field, directly in front of the LER, providing an ability to begin EVA planning on approach to each outcrop prior to egress. The vehicle can rotate 360º without any lateral movement, providing views in all directions. It has high-visibility windows, a ForeCam, AftCam, port and starboard cameras, docking cameras, and a GigaPan camera. EVA capabilities: To reduce timeline, mass, and volumetric overhead, rapid egress and ingress were envisioned, replacing an airlock with lower cabin pressure than on the International Space Station and suit ports on the aft cabin wall [2]. When needed for closer inspection and sample collecting, crew could egress in about 10 minutes through suit ports. Crew use SuitCams for additional photo-documentation, transmit mobile observations verbally, and collect surface materials. Typical simulations involved 3 to 4 EVA stations/day and 2 to 3 hr/day of boots on the ground. This allowed crew to explore a far larger territory, with more complex geological and in situ resource utilization (ISRU) features, than would a single, longer-duration EVA at one location, while also minimizing crew time in a spacesuit. Additionally, the vehicle could be driven with crew locked into the suit ports. This approach could involve a driver in the cockpit with a suited crewmember in a suit port, or the vehicle could be driven from the aft deck with both crewmembers in their suit ports. This approach was used when distances between stops were short enough that vehicle ingress and egress were less efficient than remaining in the suits and driving. Utility of suit ports: The advantages of suit ports were clearly demonstrated in those field-based trade studies. To illustrate those advantages further, consider the consequences of a SPR without suit ports at the Apollo 17 landing site. At that site, the crew's second EVA was an approximately 18 km loop conducted in a UPR, called the Lunar Roving Vehicle (LRV), in 7 hr 36 min 56 s. The traverse was composed of 5 formal stations, plus 8 additional LRV stations where crew made brief scientific stops. In a scenario involving a SPR without suit ports, crew would go EVA through an airlock and probably be limited to a single EVA per day. In that case, crew could drive the SPR ~9 km from the landing site to station 2, go EVA, and complete station 2 tasks. However, to conduct station 3 tasks, the crew would then need to walk approximately 3 km to station 3, while ground control in Houston tele-robotically drives the LER to station 3. A walk of approximately 3 km is possible, as that is what the Apollo 14 crew did before LRVs were deployed, but it is a lengthy and potentially grueling EVA. Assuming crew completes station 3 tasks, they would likely need to re-enter the SPR, ending the day's EVA, and return to the landing site. They would not be able to walk the additional distances to stations 4 and 5 (the latter being about 6 km from station 3). Thus, crew in an SPR without suit ports would require two days to accomplish the same tasks Apollo 17 crew completed in a single day. If a future crew is involved in long duration traverses on the lunar surface, the deployment of a vehicle with suit ports would probably be a better solution.

Kring, David A.↗

Mitigation of Engine Inlet Distortion Through Adjoint-Based Design

The adjoint-based design capability in FUN3D is extended to allow efficient gradient- based optimization and design of concepts with highly integrated aero-propulsive systems. A circumferential distortion calculation, along with the derivatives needed to perform adjoint-based design, have been implemented in FUN3D. This newly implemented distortion calculation can be used not only for design but also to drive the existing mesh adaptation process and reduce the error associated with the fan distortion calculation. The design capability is demonstrated by the shape optimization of an in-house aircraft concept equipped with an aft fuselage propulsor. The optimization objective is the minimization of flow distortion at the aerodynamic interface plane of this aft fuselage propulsor.

Ordaz, Irian↗

Asymmetric Thrust Reversers

An aircraft includes a propulsion supported within an aft portion of a fuselage A thrust reverser is mounted in the aft portion of the fuselage proximate the propulsion system for directing thrust in a direction to slow the aircraft. The thrust reverser includes an upper blocker door movable about a first pivot axis to a deployed position and a lower blocker door movable about a second pivot axis not parallel to the first pivot axis.

Chandler, Jesse M.↗

Aircraft System Noise Assessment of the NASA D8 Subsonic Transport Concept

Avehicle-level noise assessment has been performed for theNASAD8 concept aircraft (ND8) in the NASA Advanced Air Transport Technology Project portfolio. The NASA research-level Aircraft NOise Prediction Program (ANOPP-Research) was used to predict the noise from each source component on the ND8 to build up a noise estimate for the full aircraft. The propulsion airframe aeroacoustic (PAA) effects of the ND8, namely boundary layer ingestion (BLI) with its influence on fan noise, and the noise shielding, reflection, and diffraction mechanisms of the unconventional airframe, were empirically modeled using experimental data. Noise reduction technologies appropriate to the 2025-2035 time frame were included in this study. Including all technologies and PAA effects, the ND8 is predicted to have a cumulative margin to the Stage 4 certification metric of only 7.4 EPNdB. Boundary layer ingestion is predicted to have a detrimental impact on cumulative noise levels on the order of 15 EPNdB. Fan noise is seen to be the primary noise source at all three certification points, even if the BLI noise impact could be entirely suppressed. The impact of engine noise shielding by the airframe is limited by a lack of aft shielding and the presence of horizontal tail reflections in the aft direction. The physical constraint on engine size by the pi-tail is seen as a potential barrier to engine noise reduction through the corresponding limitation on fan bypass ratio. Mildly reduced climb performance (compared to similar reference aircraft) does not provide any benefit through increased noise propagation distance. If the boundary layer ingestion noise penalty could be suppressed such that BLI would have no effect on noise, the cumulative margin to Stage 4 would increase to 22.4 EPNdB, still below the NASA Mid Term goal of 32-42 EPNdB.

Clark, Ian A.↗

Optimizing a Coupled Solution for the Orion MPCV Using DPLR and NEQAIR

The goal of this project is to study the convective and radiative heat flux predictions using computational fluid dynamics (CFD) on a smooth body representation of the NASA Orion Multi-Purpose Crew Vehicle (MPCV). The Orion MPCV is designed for long duration missions in space and will play a crucial role in transporting astronauts in the upcoming Artemis missions. With Orion's aim for cost-effective reusability comes a need to study its Thermal Protection System (TPS), which allows for safe entry into the atmospheres of the Earth, Moon, and Mars. In order to study the effects of the TPS on test flights and better understand the aerothermal environment, CFD is a necessary tool used to create simulations of entry vehicles and calculate the heat load and heat flux. However, solutions are often performed using an uncoupled approach between the flow and the radiation solver. This leads to an overestimation of the overall heating effects because they do not consider radiative cooling. Furthermore, the effects of radiation on the aft-body is an area of research with significant uncertainty. Through the use of codes developed at NASA Ames, DPLR and NEQAIR, a coupled solution can be attained for the Orion capsule which will more accurately solve for the effects of radiation in both the fore and aft-body regions. This yields more accurate design margins for the TPS, which produces a better estimate of the thickness required. By developing an accurate coupled solution of the TPS on the Orion MPCV, planners will have greater flexibility in carrying out future missions to the Moon and beyond.

Design↗

Calibration of the 11x11-Foot Transonic Wind Tunnel at NASA Ames Research Center

A static pipe calibration characterizing the streamwise static pressure distribution was conducted at the 11-By 11-Foot Transonic Wind Tunnel at NASA Ames Research Center. This data is used to determine the local Mach number within the test section and evaluate buoyancy corrections to axial force measurements. The 60 foot long, 6 inch diameter pipe contained 444 static pressure taps spanning the test section and nozzle regions of the tunnel. The forward end of the pipe extends into the settling chamber and is held by four cables mounting to the tunnel shell, and the aft end is fixed on the institutional model support system. A hydraulic cylinder at the aft end of the pipe provides tension on the system to reduce vibration and to keep the pipe as level as possible throughout the test section. The previous calibration was improved upon by using pressure scanners with greater accuracy, ensuring a uniform pressure tube length for each tap to control pneumatic lag, optically tracking any streamwise movement of the pipe, and more tightly controlling the tunnel condition set points. Typically this calibration is conducted with the pipe on tunnel centerline and 33 inches below centerline for sting-mounted models and semi-span (i.e. floor-mounted) models respectively, however schedule demands permitted only the centerline calibration. The semi-span calibration is planned to be completed in the summer of 2020. Immediately following the static pipe calibration, a shorter, 9 foot static pipe used as the calibration check standard was installed to obtain its first post-calibration pressure dataset. This short static pipe consists of 148 static pressure taps distributed along the pipe section and one total pressure tap at the end of an ogive nose. Performing the calibration test and its check standard back-to-back allows this dataset to establish a reliable baseline for future calibration check standard testing. Over time the use of a calibration check standard offers the ability to assess the stability of the calibration through statistical process control in an efficient and costeffective manner, thereby potentially increasing the time required between full tunnel calibrations.

White, Bethany↗

Progress on the Development of a Step Height Sensor for the BOLT Flight Vehicle

This presentation will summarize progress on the development of a step height sensor for the BOLT flight vehicle. Two methods have been proposed to measure the step size in this effort.The first method involves the application of miniaturized camera embedded in the flight vehicle that views the step through a rearward-facing port (shown in Fig. 1a inset). The second method involves the application of a linear fiber optic array bundle to measure the intensity of light reflected back from the aft body step (concept shown in Fig. 1b). Preliminary testing with the miniaturized camera has been performed in the lab and shows that displacements of ≤10 μm can be measured using this method. Final testing with both methods will be performed on the test stand shown in Fig. 1a. This test stand will include a fine adjustment vertical translation stage that will be used to adjust the step height in steps of 1-2 μm. A high-magnification imaging system will also be used to verify the step height. The miniaturized camera and linear fiber optic array bundle will be mounted in one of several port holes with varying degrees of inclination relative to the top surface in the opposing aft-section slab. BOLT flight vehicle details and information on the steps can be found in Ref. [1].

BOLT↗

Azimuthal Noise Directivity of Non-Axisymmetric Jets

Far-field acoustic measurements of subsonic non-axisymmetric jets were made to carefully show the azimuthal character of the sound field. Nozzles with azimuthal features having periodicity of 2 to 6 show the progression of the azimuthal effects. Compared with predicted velocity profiles, the relationship between jet thickness and sound is demonstrated. The measurements show that even jets with very radical non-axisymmetry, such as an 8:1 rectangular nozzle, produce an axisymmetric sound field at Strouhal numbers less than 1. Significant azimuthal deviations are found for St> 1, mainly at far aft angles. The polar angle where the sound begins to be non-axisymmetric moves aft as the azimuthal order increases. The deviation is more extreme in amplitude with increasing jet speed in subsonic jets. As the azimuthal order reaches m=5 the pattern becomes more complicated, with increases and decreases in the same azimuthal plane, but with relatively small amplitudes. It is hoped that the data can be used in validating high-fidelity jet noise prediction methods.

Jet noise↗

Assessment of RANS-based Transition Models based on Experimental Data of the Common Research Model with Natural Laminar Flow

Transition models based on auxiliary transport equations augmenting the Reynolds-averaged Navier-Stokes (RANS) framework often rely upon transition correlations that were derived from a limited number of low-speed experiments and these models often fail to account for all of the relevant transition mechanisms and/or the variation in those mechanisms with respect to changes in the significant flow parameters. Available data from a recent experiment on the Common Research Model with Natural Laminar Flow (CRM-NLF) in the National Transonic Facility at the NASA Langley Research Center are used to assess the current transition modeling capability in NASA's OVERFLOW 2.3b solver for a swept wing configuration with nonzero taper and transonic cruise conditions. Specifically, the OVERFLOW solutions are used to evaluate the accuracy and robustness of the transport-equation-based transition models. Results highlight that the Spalart-Allmaras-based amplification factor transport (AFT-2017b) equation model and Menter’s shear-stress transport equation (SST2003)-based Langtry-Menter transition models (either with or without the modeling of crossflow transition) significantly underpredict the reported extent of laminar flow region over the entire span of the wing, irrespective of which instability mechanism(s) is expected to dominate the onset of the transition process. We show that the transition correlations underlying these models fail to account for the stabilizing effect of compressibility on the Tollmien-Schlichting transition, which is likely to be a major contributor to the underprediction of the laminar flow region on the CRM-NLF. The SST-2003-based Langtry-Menter model also appears to inaccurately predict the chordwise pressure variation along the majority of the wing span at all the flow conditions studied herein, due to how the turbulence intensity levels were enforced in the computations and how that was interfering with the functioning of the underlying turbulence model within the boundary layer. The AFT and Langtry-Menter models appear to be sensitive to the level of the freestream turbulence intensity, but the degree of sensitivity varies across the models.

CFD modeling↗

Global Ecosystem Dynamics Investigation (GEDI) Instrument Alignment and Test

NASA's Global Ecosystem Dynamics Investigation (GEDI) instrument was launched Dec. 5, 2018, and installed on the International Space Station 419 km from Earth. The GEDI is a Light Detection and Ranging (LIDAR) instrument; measuring the time of flight of transmitted laser beams to the Earth and back to determine altitude for geospatial mapping of forest canopy heights. The need for very dense cross track sampling for slope measurements of canopy height is accomplished by using three individual laser transmitter systems, where each beam is split into two beams by a birefringent crystal. Furthermore, one transmitter is equipped with a diffractive optical element splitting the two beams into four, for a total of 8 beams. The beams are reflected off of the features and imaged by an 800 mm diameter Receiver Telescope Assembly, composed of a Ritchey-Chretien telescope, a refractive aft optics assembly and focal plane array which collects and focuses the light from the 8 probe beams into the 8 science fibers, each with a field of view on the Earth subtending 300 mu rad. The dense cross-track sampling mandated a custom designed dual-fiber interface. The science fibers had to be aligned to the nominal, projected laser spots. The alignment was highly dependent on optimization and co-positioning of the fibers pair-wise due to mechanical constraints. This paper presents the end-to-end alignment and metrology of the full optical system from transmitter elements through receiver telescope, aft-optics, focal plane and receiver fibers performed at NASA Goddard Space Flight Center.[GRAPHICS]

Bente Eegholm↗

An Application of Flexible Multibody Simulations to Space Launch Systems Liftoff Pad Separation and Umbilical Disconnect

A flexible multibody dynamics approach is applied to the Space Launch System (SLS) liftoff Coupled Loads Analysis (CLA), enabling the inclusion of a large array of component nonlinearities with extreme computational efficiency. The nonlinearities include the cryogenic induced preloads due to the large rotations of the aft struts connecting the Core Stage (CS) to the boosters, the contact/separation and potential re-contact at the booster aft skirt to Mobile Launcher(ML) interface, contact/separation and potential re-contact at the ML/extensible columns interfaces, secondary disconnect of the CS umbilicals including the LOX and LH2 Tail Service Mast Umbilicals (TSMUs), and the disconnect of the upper stage umbilical, the Interim Cryogenic Propulsion Stage Umbilical (ICPSU). The ICPSU disconnect involves algorithms simulating the winch motors reeling lanyard ropes, the nonlinear modeling of ropes and hoses, the disconnect and capture of multiple umbilical ground plates by catch-nets (geometrically nonlinear models), and the large rotations of the ML gantry in order to track clearances between the lifting SLS vehicle and umbilicals rotating out of the way. The flexible multibody dynamics framework utilized for these simulations provided a systematic and efficient framework for adding complex nonlinearities at the system level which would have otherwise not been possible in standard CLAs or would have to be treated by separate local analyses thereby not accounting for the coupled system behavior.

Application↗

Design Exploration of a Transonic Cruise Slotted Airfoil

A knowledge-based aerodynamic design method, CDISC, has been leveraged for the computational design exploration of the cruise slotted airfoil as a drag-saving technology for single aisle, transonic transport aircraft. Aerodynamic predictions were generated using the NASA USM3D Reynolds-averaged Navier-Stokes flow solver, and laminar flow assessments were conducted using the NASA BLSTA3D boundary layer profile solver paired with the LASTRAC stability analysis and transition prediction software. The CDISC design method was used to parametrically vary several slotted-airfoil design variables to understand their impact on aerodynamic performance. In addition to establishing best practices for slotted airfoil design, it was observed that the historically noted skin-friction penalty for cruise slotted airfoils was attributed to the low-Reynolds-number boundary layer of the flap. At cruise, this skin-friction penalty negated any potential decrease in pressure drag enabled by the cruise slotted airfoil architecture. This observation led to the development of an Aft-Laminar Multi-element Airfoil concept that uses airfoil shaping to promote natural laminar flow on the flap. At cruise conditions, the concept is predicted to offer a 2.7% reduction in sectional drag relative to a fully turbulent supercritical airfoil. Off-design analyses showed that laminar flow could be maintained with near-cruise variations in angle of attack and Mach number, resulting in sustained aerodynamic efficiency improvements and a delay in the drag rise Mach number. A low-speed, high-lift analysis at takeoff conditions predicted a maximum lift coefficient of 2.4 could be achieved by use of a variable-camber leading edge on the main element and simple deflection of the flap. These computationally-predicted benefits have motivated future research toward an aft-laminar cruise slotted wing design as an incremental step toward the application of natural laminar flow technology on transonic commercial transports.

Brett R Hiller↗

Initial Regulatory and Certification Approach for the SUSAN Electrofan Concept

New aircraft concepts and associated technologies are required to meet emerging environ-mental impact goals, while still meeting established airworthiness certification standards and regulations. The SUbsonic Single Aft eNgine (SUSAN) Electrofan is one of these concepts, utilizing a single aft-mounted engine, electrified aircraft propulsion, and an emergency backup battery. This paper explores the current certification and regulatory barriers to this novel aircraft design and the potential methods for adapting certification and regulatory requirements for the SUSAN Electrofan concept. Although some certification requirements, such as the Extended Operations regulations for continued operation after the loss of an engine, can be met with only minor changes, other requirements require modifications to the regulations or the aircraft concept for compliance. Early insight to the certification approach for the SUSAN concept, while the design is still undergoing trade space exploration, allows the development of the aircraft alongside any modifications to certification standards.

SUSAN Electrofan↗

Quiet High Speed Fan II (22-inch) Duct Mode Characteristics as Measured by the Rotating Rake Mode Measurement System while Operated in the NASA Glenn 9x15 Low Speed Wind Tunnel

Wind Tunnel test at a tunnel Mach number of 0.10. This was an entry to investigate the effect of “stator clocking” on noise. The fan consisted of a moderately aft swept rotor and an aft swept set of stator vanes. The fan stage consisted of 22 rotor blades, 50 stator vanes, and 10 downstream support struts. A set of stator vanes designed for lower noise was tested as well as a baseline stator vane set. The stator assembly could be rotated several degrees to adjust the clocking angle between the stator vane pack and the strut assembly. All configurations were with a hard wall duct (no acoustic treatment). The NASA Glenn Research Center’s Rotating Rake Mode Measurement System was utilized to obtain a complete map of the acoustic duct modes present in the ducted fan. The system is a radial rake emersed into the duct that continuously rotates about the duct centerline. For the two stator configurations, data were acquired at several different fan speeds which included nominal, approach, cutback, and takeoff conditions. Analysis of the mode power level results at the fan fundamentals showed the improved designed set resulted in lower rotor-stator and rotor-strut interaction acoustic levels for the interaction modes. Varying the angle between the stators and struts was shown to be a viable method to achieve a minimum in rotor-strut interaction mode power level. Multiple-pure-tones generated by the Quiet High Speed Fan II in the inlet were also measured.

Turbofan, Duct Modes↗