Search NASA⌕ Search

SEARCH · Search NASA

Results for “conditional approach”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Active Flap Control of the SMART Rotor for Vibration Reduction

Active control methodologies were applied to a full-scale active flap rotor obtained during a joint Boeing/ DARPA/NASA/Army test in the Air Force National Full-Scale Aerodynamic Complex 40- by 80-foot anechoic wind tunnel. The active flap rotor is a full-scale MD 900 helicopter main rotor with each of its five blades modified to include an on-blade piezoelectric actuator-driven flap with a span of 18% of radius, 25% of chord, and located at 83% radius. Vibration control demonstrated the potential of active flaps for effective control of vibratory loads, especially normal force loads. Active control of normal force vibratory loads using active flaps and a continuous-time higher harmonic control algorithm was very effective, reducing harmonic (1-5P) normal force vibratory loads by 95% in both cruise and approach conditions. Control of vibratory roll and pitch moments was also demonstrated, although moment control was less effective than normal force control. Finally, active control was used to precisely control blade flap position for correlation with pretest predictions of rotor aeroacoustics. Flap displacements were commanded to follow specific harmonic profiles of 2 deg or more in amplitude, and the flap deflection errors obtained were less than 0.2 deg r.m.s.

Hall, Steven R.↗
NASA NTRS2012

Acoustic Performance of an Advanced Model Turbofan in Three Aeroacoustic Test Facilities

A model advanced turbofan was acoustically tested in the NASA Glenn 9- by 15-Foot-Low-Speed Wind Tunnel (LSWT), and in two other aeroacoustic facilities. The Universal Propulsion Simulator (UPS) fan was designed and manufactured by the General Electric Aircraft Engines (GEAE) Company, and featured active core, as well as bypass, flow paths. The reference test configurations were with the metal, M4, rotor with hardwall and treated bypass flow ducts. The UPS fan was tested within an airflow at a Mach number of 0.20 (limited flow data were also acquired at a Mach number of 0.25) which is representative of aircraft takeoff and approach conditions. Comparisons were made between data acquired within the airflow (9x15 LSWT and German-Dutch Wind Tunnel (DNW)) and outside of a free jet (Boeing Low Speed Aero acoustic Facility (LSAF) and DNW). Sideline data were acquired on an 89-in. (nominal 4 fan diameters) sideline using the same microphone assembly and holder in the 9x15 LSWT and DNW facilities. These data showed good agreement for similar UPS operating conditions and configurations. Distortion of fan spectra tonal content through a free jet shear layer was documented, suggesting that in-flow acoustic measurements are required for comprehensive fan noise diagnostics. However, there was good agreement for overall sound power level (PWL) fan noise measurements made both within and outside of the test facility airflow.

Woodward, Richard P.↗
NASA NTRS2012

An Electrostatic Precipitator System for the Martian Environment

Human exploration missions to Mars will require the development of technologies for the utilization of the planet's own resources for the production of commodities. However, the Martian atmosphere contains large amounts of dust. The extraction of commodities from this atmosphere requires prior removal of this dust. We report on our development of an electrostatic precipitator able to collect Martian simulated dust particles in atmospheric conditions approaching those of Mars. Extensive experiments with an initial prototype in a simulated Martian atmosphere showed efficiencies of 99%. The design of a second prototype with aerosolized Martian simulated dust in a flow-through is described. Keywords: Space applications, electrostatic precipitator, particle control, particle charging

Calle, C. I.↗
NASA NTRS2015

Assessment of Slat Noise Predictions for 30P30N High-Lift Configuration From BANC-III Workshop

This paper presents a summary of the computational predictions and measurement data contributed to Category 7 of the 3rd AIAA Workshop on Benchmark Problems for Airframe Noise Computations (BANC-III), which was held in Atlanta, GA, on June 14-15, 2014. Category 7 represents the first slat-noise configuration to be investigated under the BANC series of workshops, namely, the 30P30N two-dimensional high-lift model (with a slat contour that was slightly modified to enable unsteady pressure measurements) at an angle of attack that is relevant to approach conditions. Originally developed for a CFD challenge workshop to assess computational fluid dynamics techniques for steady high-lift predictions, the 30P30N configurations has provided a valuable opportunity for the airframe noise community to collectively assess and advance the computational and experimental techniques for slat noise. The contributed solutions are compared with each other as well as with the initial measurements that became available just prior to the BANC-III Workshop. Specific features of a number of computational solutions on the finer grids compare reasonably well with the initial measurements from FSU and JAXA facilities and/or with each other. However, no single solution (or a subset of solutions) could be identified as clearly superior to the remaining solutions. Grid sensitivity studies presented by multiple BANC-III participants demonstrated a relatively consistent trend of reduced surface pressure fluctuations, higher levels of turbulent kinetic energy in the flow, and lower levels of both narrow band peaks and the broadband component of unsteady pressure spectra in the nearfield and farfield. The lessons learned from the BANC-III contributions have been used to identify improvements to the problem statement for future Category-7 investigations.

Choudhari, Meelan↗
NASA NTRS2015

On the Connection Between Flap Side-Edge Noise and Tip Vortex Dynamics

The goal of the present work is to investigate how the dynamics of the vortical flow about the flap side edge of an aircraft determine the acoustic radiation. A validated lattice- Boltzmann CFD solution of the unsteady flow about a detailed business jet configuration in approach conditions is used for the present analysis. Evidence of the connection between the noise generated by several segments of the inboard flap tip and the aerodynamic forces acting on the same segments is given, proving that the noise generation mechanism has a spatially coherent and acoustically compact character on the scale of the flap chord, and that the edge-scattering effects are of secondary importance. Subsequently, evidence of the connection between the kinematics of the tip vortex system and the aerodynamic force is provided. The kinematics of the dual vortex system are investigated via a core detection technique. Emphasis is placed on the mutual induction effects between the two main vortices rolling up from the pressure and suction sides of the flap edge. A simple heuristic formula that relates the far-field noise spectrum and the cross-spectrum of the unsteady vortical positions is developed.

Casalino, D.↗
NASA NTRS2017

An Assessment of the Icing Blade and the SEA Multi-Element Sensor for Liquid Water Content Calibration of the NASA GRC Icing Research Tunnel

The Icing Research Tunnel at NASA Glenn has recently switched to from using the Icing Blade to using the SEA Multi-Element Sensor (also known as the multi-wire) for its calibration of cloud liquid water content. In order to perform this transition, tests were completed to compare the Multi-Element Sensor to the Icing Blade, particularly with respect to liquid water content, airspeed, and drop size. The two instruments were found to compare well for the majority of Appendix C conditions. However, it was discovered that the Icing Blade under-measures when the conditions approach the Ludlam Limit. This paper also describes data processing procedures for the Multi-Element Sensor in the IRT, including collection efficiency corrections, mounting underneath a splitter plate, and correcting for a jump in the compensation wire power. Further data is presented to describe the repeatability of the IRT with the Multi-Element sensor, health-monitoring checks for the instrument, and a sensing-element configuration comparison.

multi-element sensor↗
NASA NTRS2016

An Assessment of the Icing Blade and the SEA Multi-Element Sensor for Liquid Water Content Calibration of the NASA GRC Icing Research Tunnel

The Icing Research Tunnel at NASA Glenn has recently switched from using the Icing Blade to using the SEA Multi-Element Sensor (also known as the multi-wire) for its calibration of cloud liquid water content. In order to peform this transition, tests were completed to compare the Multi-Element Sensor to the Icing Blade, particularly with respect to liquid water content, airspeed, and drop size. The two instruments were found to compare well for the majority of Appendix C conditions. However, it was discovered that the Icing Blade under-measures when the conditions approach the Ludlam Limit. This paper also describes data processing procedures for the Multi-Element Sensor in the IRT, including collision efficiency corrections, mounting underneath a splitter plate, and correcting for a jump in the compensation wire power. Further data is presented to describe the repeatability of the IRT with the Multi-Element Sensor, health-monitoring checks for the instrument, and a sensing-element configuration comparison. Ultimately these tests showed that in the IRT, the multi-wire is a better instrument for measuring cloud liquid water content than the blade.

aircraft icing↗
NASA NTRS2017

Approach Considerations in Aircraft with High-Lift Propeller Systems

NASA's research into distributed electric propulsion (DEP) includes the design and development of the X-57 Maxwell aircraft. This aircraft has two distinct types of DEP: wingtip propellers and high-lift propellers. This paper focuses on the unique opportunities and challenges that the high-lift propellers--i.e., the small diameter propellers distributed upstream of the wing leading edge to augment lift at low speeds--bring to the aircraft performance in approach conditions. Recent changes to the regulations related to certifying small aircraft (14 CFR x23) and these new regulations' implications on the certification of aircraft with high-lift propellers are discussed. Recommendations about control systems for high-lift propeller systems are made, and performance estimates for the X-57 aircraft with high-lift propellers operating are presented.

Patterson, Michael D.↗
NASA NTRSDate not supplied

Auxiliary Inlet Design Study for Mach 1.4

Auxiliary inlet configurations were developed for inlets designed for a NASA Mach 1.4 Supersonic Technology Concept Aeroplane (STCA). The inlets included axisymmetric pitot and axisymmetric spike inlets. The auxiliary inlets were full circumferential slots within the cowl. Computational fluid dynamics (CFD) simulations were performed to obtain inlet performance metrics of total pressure recovery and radial distortion at the engine face for take-off and approach conditions. Methods of design of experiments were used to explore the statistical significance of the design factors for the auxiliary inlets which included the axial location, length of the opening, and auxiliary inlet angle. The results demonstrate the effect of the design factors and show the potential for improved inlet performance with the use of auxiliary inlets.

supersonic inlets↗
NASA NTRSDate not supplied

Prediction of Quadrotor Acoustics Using RVLT Toolchain

The NASA Revolutionary Vertical Lift Technology (RVLT) Project toolchain is used to predict the acoustics of the NASA 6-passenger quadrotor concept vehicle. A qualitative study of the blade vertical loading to understand noise trends with tip speed and number of blades was performed. Trends of the individual rotor noise sources (thickness, loading, and broadband) are studied. Noise is predicted and analyzed for three flight conditions: approach (6-deg descent), level flight flyover, and takeoff. Three quadrotor designs are considered: tip speed=700 ft/sec, 3 blades (“700/3”), “550/3”, and “375/7”. The qualitative study shows that during approach multiple, strong BVIs occur on the advancing and retreating sides for “700/3” and “550/3”, but for “375/7” only weak BVIs are present. Consistent with this BVI scenario, the quadrotor noise results show that “700/3” has the highest loading noise (and EPNL) and “375/7” the lowest loading noise (and EPNL), with “550/3” falling in between. Broadband noise during approach and flyover is roughly the same for all three designs. Takeoff broadband noise is much lower, especially for “375/7”. The bulk of the results have been obtained with a rigid uniform blade model, with a flap hinge and a pitch bearing. For “550/3”, noise predictions with elastic nonuniform blades have been initiated and are ongoing.

Quadrotor↗
NASA NTRSDate not supplied

Effects of Porous Gap Fillers on Leading-Edge Slat Noise of 30P30N. Part II: PIV Measurements

The leading edge slat of a high-lift system is one of the main contributors to the airframe noise during approach conditions. This paper, the second of two parts, continues our previous studies on the slat gap filler as a passive noise control device on the two-dimensional, 30P30N multielement airfoil. Whereas the earlier study was focused on the effects of an impermeable gap filler that completely blocks the flow through the gap, this follow-on assessment is devoted to permeable slat gap fillers that allow limited amounts of flow to pass through the gap. Part I of this two-part abstract described the aerodynamic and acoustic effects of the permeable gap fillers, as inferred from both the measurements of static and unsteady surface pressures and the microphone array data for the radiated noise. To understand the physical mechanism responsible for the noise reduction documented in Part I, as well as for the accompanying aerodynamic penalty due to the porous gap fillers, Particle Image Velocimetry (PIV) is used in this paper to measure the flow details in the slat-cove region of the 30P30N model. A single angle of attack (𝛼 = 5.5◦) and a chord based Reynolds number of 1.71𝑒6 are selected as the test conditions. The PIV results show that the slat flow features are significantly altered with the presence of the porous gap filler, resulting in a more stable slat-cove shear layer and, thus, reduced velocity fluctuations with a successive decrease in the permeability. The porous gap filler with the lowest permeability acts similar to the solid gap filler. However, flow separation is observed on the upper side of the porous interface, which leads to an aerodynamic performance penalty via a reduction in lift on the main wing.

slat noise↗
NASA NTRSDate not supplied

Supercritical Water Oxidation: Testing of Ersatz Wastewater

Supercritical Water Oxidation (SCWO) technology is under consideration by NASA for treatment of wastewater and other wet waste streams. The dramatic changes in water's thermophysical properties near its critical point (374◦C and 22 MPa) result in nearly complete solubility of organics and gases. The absence of inter–phase transport processes and phase separation results in dramatically reduced reaction timescales and the elimination of complications of two–phase transport and processing in reduced gravity. Other advantages are that the product streams are microbially inert,benign, amenable to resource recovery and organic waste. This paper first described the design and construction of tubular SCWO reactor built at NASA Glenn. It consists of separate wastewater and oxidizer inlets, which are independently heated and pressurized to conditions approaching the critical point of water. These flows enter a reactor chamber having a volume of 62 ml. The reactor section is heated to the desired temperature in the range of 450 ◦C – 650 ◦C. A pressure regulator maintains the operating pressure nominally between 27.6 MPa and 28.3 MPa. The reacted fluid is captured downstream of the pressure regulator during an experiment run. A separate vent line may also be used for collection of the effluent. Next, the results of SCWO experiments using both a representative organic (ethanol) as well as a wastewater simulant provided by NASA Ames are presented. Raman analysis of the pre–reacted and post–reacted solutions are presented. Finally, initial efforts on supplementary chemical kinetics modeling of ethanol oxidation with and without inclusion of a nitrogenous compounds (ammonia) are presented.

Michael C Hicks↗
NASA NTRSDate not supplied

Fan Noise Predictions of the NASA Source Diagnostic Test using Unsteady Simulations with LAVA Part II - Tonal and Broadband Noise Assessment

The NASA Source Diagnostic Test (SDT) campaign experimental data is used for validation of a sliding mesh technique recently implemented within the Launch, Ascent, and Vehicle Aerodynamics computational framework for time-accurate simulation of rotating fans. The far-field acoustics are analyzed in this work, building upon the aerodynamic validation studies previously published in Part I. Two modeling approaches are explored: the unsteady Reynolds-averaged Navier Stokes (uRANS) with the negative Spalart-Allmaras (SA-neg) turbulence model closure, and a hybrid Reynolds-averaged Navier Stokes/large-eddy simulation (RANS/LES) paradigm employing a zonal detached-eddy simulation (ZDES) closure with enhanced shielding protection. Two convective flux scheme approaches with different dissipation properties are also explored with ZDES. The Ffowcs-Williams and Hawkings (FW-H) permeable surface approach is used for propagation of the near-field acoustics to the far-field microphone locations. This work analyzes the low-speed (approach) condition, characterized by a fan rotation speed of 7808 rotations-per-minute (RPM). Three different grid levels ranging between 200 million and 1.1 billion grid points are considered. Results show good prediction of the broadband noise levels at sideline angles ranging between 70° and 110°. The forward arc observers show an under-prediction of the overall sound pressure levels (OASPL) even at the fine grid level. A breakdown of the inlet and exhaust contributions reveals a steep drop-off in the broadband noise levels past a blade-passing frequency (BPF) of 1.5, potentially caused by a lack of resolved small-scale turbulent fluctuations in the interstage region. Past 110° the OASPL are over-predicted by up to 10 dB due to an over-prediction of the low-frequency broadband noise levels in the aft arc. The detuning of BPF2 caused by small deviations in the blade stagger angle around the wheel is captured, and a corresponding decrease in the sound power level for this cut-on tone is observed.

AATT↗
NASA NTRSDate not supplied

Flametube Evaluation of a Lean-Lean Combustor Concept Developed for Supersonic Cruise Aircraft

Gaseous emissions were measured in single-cup flametube tests of an advanced low-NOx combustor concept at simulated supersonic cruise conditions. The combustor concept is a low technology readiness level (TRL), lean front-end design developed under the NASA Fundamental Aeronautics/Supersonics project to minimize NOx emissions at supersonic cruise. The flametube conditions matched or approached combustor conditions at supersonic cruise, with combustor inlet temperatures up to 920 K, inlet pressures up to 19 bar, and combusted gas temperatures up to 2,120 K. Whether these conditions met or just approached supersonic cruise conditions depended on the type of engine the combustor would be installed in. Two types of engines were considered here: a “derivative” engine based on a current technology and an “advanced” engine with a higher operating pressure ratio and higher temperature limits. For the “derivative” engine, the combustor is expected to be at least close to meeting the NASA NOx emissions goal of 10 g-NOx/kg-fuel at supersonic cruise. However, with the higher combustor inlet and flame temperatures of the advanced engine, NOx emissions are expected to be well above the goal.

combustion↗
NASA NTRS2022

Flametube Evaluation of a Lean-Lean Combustor Concept Developed for Supersonic Cruise Aircraft

Gaseous emissions were measured in single-cup flametube tests of an advanced low-NO x combustor concept at simulated supersonic cruise conditions. The combustor concept is a low technology readiness level (TRL), lean front-end design developed under the NASA Fundamental Aeronautics/Supersonics project to minimize NO x emissions at supersonic cruise. The flametube conditions matched or approached combustor conditions at supersonic cruise, with combustor inlet temperatures up to 920 K, inlet pressures up to 19 bar, and combusted gas temperatures up to 2,120 K. Whether these conditions met or just approached supersonic cruise conditions depended on the type of engine the combustor would be installed in. Two types of engines were considered here: a "derivative" engine based on a current technology and an "advanced" engine with a higher operating pressure ratio and higher temperature limits. For the "derivative" engine, the combustor is expected to be at least close to meeting the NASA NO x emissions goal of 10 g-NO x /kg-fuel at supersonic cruise. However, with the higher combustor inlet and flame temperatures of the advanced engine, NO x emissions are expected to be well above the goal.

supersonic cruise↗
NASA NTRS1986

Multiple boundary condition test (MBCT) approach to update mathematical models of large flexible structures

A major challenge to the structural dynamicist is to validate mathematical models of large space structures which cannot be ground tested because of its size and/or flexibility. The paper presents a Multiple Boundary Condition Test (MBCT) approach which allows a systematic validation of the mathematical model by performing a number of ground tests on a large structure with variable boundary conditions. A numerical simulation is presented which illustrates the validity of the MBCT including some of the potential limitations.

Wada, B. K.↗
NASA NTRS1992

Stability margins for multilinear interval systems by way of phase conditions: A unified approach

A simple way of checking the stability with respect to an arbitrary stability region of a family of polynomials containing a vector of parameters varying within prescribed intervals is discussed. It is assumed that the parameters appear affine multilinearly in the characteristic polynomial coefficients. The condition proposed is simply to check the phase difference of the vertex polynomials. This test based on the mapping theorem significantly reduces computational complexity. Mathematical proofs are omitted. The results can be used to determine various stability margins of control systems containing interconnected interval subsystems. These include the gain, phase, time-delay, H(sup infinity), and nonlinear sector bounded stability margins of multilinear interval systems.

Keel, L. H.↗