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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.

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At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Helicopter tail rotor noise

A study was made of helicopter tail rotor noise, particularly that due to interactions with the main rotor tip vortices, and with the fuselage separation mean wake. The tail rotor blade-main rotor tip vortex interaction is modelled as an airfoil of infinite span cutting through a moving vortex. The vortex and the geometry information required by the analyses are obtained through a free wake geometry analysis of the main rotor. The acoustic pressure-time histories for the tail rotor blade-vortex interactions are then calculated. These acoustic results are compared to tail rotor loading and thickness noise, and are found to be significant to the overall tail rotor noise generation. Under most helicopter operating conditions, large acoustic pressure fluctuations can be generated due to a series of skewed main rotor tip vortices passing through the tail rotor disk. The noise generation depends strongly upon the helicopter operating conditions and the location of the tail rotor relative to the main rotor.

Chou, S.-T.↗
NASA NTRS1987

Helicopter tail rotor blade-vortex interaction noise

A study is made of helicopter tail rotor noise, particularly that due to the interactions with main rotor tip vortices. Summarized here are present analysis, the computer codes, and the results of several test cases. Amiet's unsteady thin airfoil theory is used to calculate the acoustics of blade-vortex interaction. The noise source is modelled as a force dipole resulting from an airfoil of infinite span chopping through a skewed line vortex. To analyze the interactions between helicopter tail rotor and main rotor tip vortices, we developed a two-step approach: (1) the main rotor tip vortex system is obtained through a free wake geometry calculation of the main rotor using CAMRAD code; (2) acoustic analysis takes the results from the aerodynamic interaction analysis and calculates the farfield pressure signatures for the interactions. It is found that under a wide range of helicopter flight conditions, acoustic pressure fluctuations of significant magnitude can be generated by tail rotors due to a series of interactions with main rotor tip vortices. This noise mechanism depends strongly on the helicopter flight conditions and the relative location and phasing of the main and tail rotors. fluctuations of significant magnitude can be generated by tail rotors due to a series of interactions with main rotor tip vortices. This noise mechanism depends strongly upon the helicopter flight conditions and the relative location and phasing of the main and tail rotors.

George, Albert R.↗
NASA NTRS1994

Study of noise on a small-scale hovering tilt rotor

A study of the noise from a small-scale, semi-span test of a hovering tilt rotor is presented. The effect on noise of the tilt rotor fountain and of several tilt rotor configuration changes is examined. Measurments were made of an isolated rotor; rotor and wing; rotor and image plane; and rotor, wing and image plane. This last configuration models a tilt rotor. With a tilt rotor, the wing turns some of the rotor downwash inboard. When the flows meet, some of the flow is forced upward creating a fountain. Some of the fountain is ingested into the rotor disc. Rotor fountain interactions produce very unsteady impulsive noise to the rear of the aircraft. Configuration changes to investigate download reduction included increasing the rotor/wing separation, deflecting the flap, adding the nacelle, adding vortex trapping plates and blowing on the upper surface. These configuration changes produced little or no effect on the noise, although they changed the net download. Increasing the wing span, however, reduced the annoying content of this impulse noise about 2dB.

Mosher, Marianne↗
NASA NTRS1996

Full-scale S-76 rotor performance and loads at low speeds in the NASA Ames 80- by 120-Foot Wind Tunnel

A full-scale helicopter rotor test was conducted in the NASA Ames 80- by 120-Foot Wind Tunnel with a four-bladed S-76 rotor system. Rotor performance and loads data were obtained over a wide range of rotor shaft angles-of-attack and thrust conditions at tunnel speeds ranging from 0 to 100 kt. The primary objectives of this test were (1) to acquire forward flight rotor performance and loads data for comparison with analytical results; (2) to acquire S-76 forward flight rotor performance data in the 80- by 120-Foot Wind Tunnel to compare with existing full-scale 40- by 80-Foot Wind Tunnel test data that were acquired in 1977; (3) to evaluate the acoustic capability of the 80- by 120- Foot Wind Tunnel for acquiring blade vortex interaction (BVI) noise in the low speed range and compare BVI noise with in-flight test data; and (4) to evaluate the capability of the 80- by 120-Foot Wind Tunnel test section as a hover facility. The secondary objectives were (1) to evaluate rotor inflow and wake effects (variations in tunnel speed, shaft angle, and thrust condition) on wind tunnel test section wall and floor pressures; (2) to establish the criteria for the definition of flow breakdown (condition where wall corrections are no longer valid) for this size rotor and wind tunnel cross-sectional area; and (3) to evaluate the wide-field shadowgraph technique for visualizing full-scale rotor wakes. This data base of rotor performance and loads can be used for analytical and experimental comparison studies for full-scale, four-bladed, fully articulated rotor systems. Rotor performance and structural loads data are presented in this report.

Shinoda, Patrick M.↗
NASA NTRS1996

Wave rotor demonstrator engine assessment

The objective of the program was to determine a wave rotor demonstrator engine concept using the Allison 250 series engine. The results of the NASA LERC wave rotor effort were used as a basis for the wave rotor design. A wave rotor topped gas turbine engine was identified which incorporates five basic requirements of a successful demonstrator engine. Predicted performance maps of the wave rotor cycle were used along with maps of existing gas turbine hardware in a design point study. The effects of wave rotor topping on the engine cycle and the subsequent need to rematch compressor and turbine sections in the topped engine were addressed. Comparison of performance of the resulting engine is made on the basis of wave rotor topped engine versus an appropriate baseline engine using common shaft compressor hardware. The topped engine design clearly demonstrates an impressive improvement in shaft horsepower (+11.4%) and SFC (-22%). Off design part power engine performance for the wave rotor topped engine was similarly improved including that at engine idle conditions. Operation of the engine at off design was closely examined with wave rotor operation at less than design burner outlet temperatures and rotor speeds. Challenges identified in the development of a demonstrator engine are discussed. A preliminary design was made of the demonstrator engine including wave rotor to engine transition ducts. Program cost and schedule for a wave rotor demonstrator engine fabrication and test program were developed.

Snyder, Philip H.↗
NASA NTRS2002

Flywheel Rotor Safe-Life Technology

Since the 1960s, research has been conducted into the use of flywheels as energy storage systems. The-proposed applications include energy storage for hybrid and electric automobiles, attitude control and energy storage for satellites, and uninterruptible power supplies for hospitals and computer centers. For many years, however, the use of flywheels for space applications was restricted by the total weight of a system employing a metal rotor. With recent technological advances in the manufacturing of composite materials, however, lightweight composite rotors have begun to be proposed for such applications. Flywheels with composite rotors provide much higher power and energy storage capabilities than conventional chemical batteries. However, the failure of a high speed flywheel rotor could be a catastrophic event. For this reason, flywheel rotors are classified by the NASA Fracture Control Requirements Standard as fracture critical parts. Currently, there is no industry standard to certify a composite rotor for safe and reliable operation forth( required lifetime of the flywheel. Technical problems hindering the development of this standard include composite manufacturing inconsistencies, insufficient nondestructive evaluation (NDE) techniques for detecting defects and/or impact damage, lack of standard material test methods for characterizing composite rotor design allowables, and no unified proof (over-spin) test for flight rotors. As part of a flywheel rotor safe-life certification pro-ram funded b the government, a review of the state of the art in composite rotors is in progress. The goal of the review is to provide a clear picture of composite flywheel rotor technologies. The literature review has concentrated on the following topics concerning composites and composite rotors: durability (fatigue) and damage tolerance (safe-life) analysis/test methods, in-service NDE and health monitoring techniques, spin test methods/ procedures, and containment options. This report presents the papers selected for their relevance to this topic and summarizes them.

Ratner, J. K. H.↗
NASA NTRS2013

FUN3D Airload Predictions for the Full-Scale UH-60A Airloads Rotor in a Wind Tunnel

An unsteady Reynolds-Averaged Navier-Stokes solver for unstructured grids, FUN3D, is used to compute the rotor performance and airloads of the UH-60A Airloads Rotor in the National Full-Scale Aerodynamic Complex (NFAC) 40- by 80-foot Wind Tunnel. The flow solver is loosely coupled to a rotorcraft comprehensive code, CAMRAD-II, to account for trim and aeroelastic deflections. Computations are made for the 1-g level flight speed-sweep test conditions with the airloads rotor installed on the NFAC Large Rotor Test Apparatus (LRTA) and in the 40- by 80-ft wind tunnel to determine the influence of the test stand and wind-tunnel walls on the rotor performance and airloads. Detailed comparisons are made between the results of the CFD/CSD simulations and the wind tunnel measurements. The computed trends in solidity-weighted propulsive force and power coefficient match the experimental trends over the range of advance ratios and are comparable to previously published results. Rotor performance and sectional airloads show little sensitivity to the modeling of the wind-tunnel walls, which indicates that the rotor shaft-angle correction adequately compensates for the wall influence up to an advance ratio of 0.37. Sensitivity of the rotor performance and sectional airloads to the modeling of the rotor with the LRTA body/hub increases with advance ratio. The inclusion of the LRTA in the simulation slightly improves the comparison of rotor propulsive force between the computation and wind tunnel data but does not resolve the difference in the rotor power predictions at mu = 0.37. Despite a more precise knowledge of the rotor trim loads and flight condition, the level of comparison between the computed and measured sectional airloads/pressures at an advance ratio of 0.37 is comparable to the results previously published for the high-speed flight test condition.

Lee-Rausch, Elizabeth M.↗
NASA NTRSDate not supplied

ELISA: A Tool for Optimization of Rotor Hover Performance at Low Reynolds Number in the Mars Atmosphere

The Evolutionary aLgorithm for Iterative Studies of Aeromechanics (ELISA) was developed in support of the Rotorcraft Optimization for the Advancement of Mars eXploration (ROAMX) project. ELISA was developed to enable aerodynamic rotor hover optimization for low Reynolds number flows in the Mars atmosphere. The first objective of the algorithm allows for unconventional airfoil parameterization and multi-objective airfoil geometry optimization using OVERFLOW. The Pareto-optimal airfoil sets are converted to a set of Pareto-optimal airfoil decks, providing the lowest drag air foil geometry for each angle of attack, removing the need to arbitrarily select the airfoils to be used in the rotor optimization. The second objective allows for rotor geometry optimization with simultaneous maximization of blade loading and minimization of rotor power using the comprehensive analysis CAMRADII. The result is a Pareto-optimal rotor set, providing the lowest power rotor for each attainable blade loading, and one of the first tools for hover-optimized rotors for high-subsonic low Reynolds number conditions. The airfoil thickness can be modified after the airfoil optimization is complete, allowing for a post-airfoil-optimization adjustment of blade thickness to facilitate conforming to external structural analyses requirements. The relevance of the code is demonstrated with case studies for the ROAMX rotor optimization for Ingenuity-sized single rotors in the Mars atmosphere, a performance study optimizing the chord and twist of Ingenuity’s coaxial rotor resulting in the Sample Recovery Helicopters candidate rotor, and high-subsonic low Reynolds number airfoil optimization providing novel insights for higher-efficiency low Reynolds number airfoil geometries and flow physics.

ELISA↗
NASA NTRS1981

Comparison of upwind and downwind rotor operation of the DOE/NASA 100-kW MOD-0 wind turbine

Tests were conducted on a 38m diameter horizontal axis wind turbine, which had first a rotor downwind of the supporting truss tower and then upwind of the tower. Aside from the placement of the rotor and the direction of rotation of the drive train, the wind turbine was identical for both tests. Three aspects of the test results are compared: rotor blade bending loads, rotor teeter response, and nacelle yaw moments. As a result of the tests, it is shown that while mean flatwise bending moments were unaffected by the placement of the rotor, cyclic flatwise bending tended to increase with wind speed for the downwind rotor while remaining somewhat uniform with wind speed for the upwind rotor, reflecting the effects of increased flow disturbance for downwind rotor. Rotor teeter response was not significantly affected by the rotor location relative to the tower, but appears to reflect reduced teeter stability near rated wind speed for both configurations. Teeter stability appears to return above rated wind speed, however. Nacelle yaw moments are higher for the upwind rotor but do not indicate significant design problems for either configuration.

Glasgow, J. C.↗
NASA NTRS1981

General purpose research rotor

An analytical study, under a NASA contract, is performed on an advanced flight research rotor (four-bladed, 54 ft in diameter, with bearingless rotor retention characteristics) to determine the sensitivity of total rotor characteristics such as vibratory hub loads, rotor horsepower, and blade loads, to parametric variability of the rotor configuration. The sensitivity of the rotor to various combinations of blade planform taper, percent of blade span that is tapered, tip sweep angle, built-in-twist, and torsional frequency is determined for specific configurations by randomly selecting combinations of these parameters. Characteristics of other intermediate rotor configurations are determined by a regression analysis. The results show that a wide range of rotor total performance characteristics can be obtained for a rotor trimmed to the same flight conditions. The regression equations predict total performance of the rotor very well and appear to be a useful analytical tool for rotor design optimization. Figures showing the results of the various tests are given along with a table of the regression coefficients.

Jones, R.↗
NASA NTRS1984

Numerical analysis of a variable camber rotor blade as a lift control device

A new rotor configuration called the variable camber rotor was numerically investigated as a lift control device. This rotor differs from a conventional (baseline) rotor only in the blade aft section. In this configuration, the aft section or flap is attached to the forward section by pin joint arrangement, and also connected to the rotor control system for the control of rotor thrust level and vectoring. Pilot action to the flap deflection controls rotor lift and tip path plane tilt. The drag due to flaps is presented and the theoretical result correlated with test data. The assessment of payoff for the variable camber rotor in comparison with conventional (baseline) rotor was examined in hover. The variable camber rotor is shown to increase hover power required by 1.35%, but such a minimal power penalty is not significant enough to be considered a negative result. In forward flight, the control needs of the variable camber rotor were evaluated.

Awani, A. O.↗
NASA NTRS1984

Hover test of a full-scale hingeless rotor

The performance and aeroelastic stability in hover of a 9.8-m diameter, hingeless helicopter rotor system was evaluated. Rotor performance and inplane damping data were obtained for rotor operation between 350 and 425 rpm for thrust coefficients (CT/sigma) between 0.0 and 0.12. At constant rotor thrust, a minimum in rotor inplane damping was measured at 400 rpm. Good agreement is shown between experimental performance data and predicted performance. The influence of different aerodynamic inflow models on predicting damping levels is also shown. The best correlation with experimental stability data was obtained when a dynamic inflow model was used instead of static or quasistatic inflow models. Comparison with other full scale, hingeless rotor data in hover is presented. The hingeless rotor data and data from a full scale, bearingless main rotor test performed on the same general purpose test apparatus were compared. Although the bearingless rotor was more highly damped at design tip speed and 1-g thrust operation, greater sensitivity to operating conditions is shown. At low thrust levels the bearingless main rotor is less damped than the hingeless rotor.

Warmbrodt, W.↗
NASA NTRS1984

Hover test of a full-scale hingeless rotor

The performance and aeroelastic stability in hover of a 9.8-m diameter, hingeless helicopter rotor system was evaluated. Rotor performance and inplane damping data were obtained for rotor operation beteen 350 and 425 rpm for thrust coefficients (CT/sigma) between 0.0 and 0.12. At constant rotor thrust, a minimum in rotor inplane damping was measured at 400 rpm. Good agreement is shown between experimental performance data and predicted performance. The influence of different aerodynamic inflow models on predicting damping levels is also shown. The best correlation with experimental stability data was obtained when a dynamic inflow model was used instead of static or quasistatic inflow models. Comparison with other full scale, hingeless rotor data in hover is presented. The hingeless rotor data and data from a full scale, bearingless main rotor test performed on the same general purpose test apparatus were compared. Although the bearingless rotor was more highly damped at design tip speed and 1-g thrust operation, greater sensitivity to operating conditions is shown. At low thrust levels the bearingless main rotor is less damped than the hingeless rotor.

Warmbrodt, W.↗
NASA NTRS1993

Navier-Stokes flowfield computation of wing/rotor interaction for a tilt rotor aircraft in hover

The download on the wing produced by the rotor-induced downwash of a tilt rotor aircraft in hover is of major concern because of its severe impact on payload-carrying capability. A method has been developed to help gain a better understanding of the fundamental fluid dynamics that causes this download, and to help find ways to reduce it. In particular, the method is employed in this work to analyze the effect of a tangential leading edge circulation-control jet on download reduction. Because of the complexities associated with modeling the complete configuration, this work focuses specifically on the wing/rotor interaction of a tilt rotor aircraft in hover. The three-dimensional, unsteady, thin-layer compressible Navier-Stokes equations are solved using a time-accurate, implicit, finite difference scheme that employs LU-ADI factorization. The rotor is modeled as an actuator disk which imparts both a radical and an azimuthal distribution of pressure rise and swirl to the flowfield. A momentum theory blade element analysis of the rotor is incorporated into the Navier-Stokes solution method. Solution blanking at interior points of the mesh has been shown here to be an effective technique in introducing the effects of the rotor and tangential leading edge jet. Results are presented both for a rotor alone and for wing/rotor interaction. The overall mean characteristics of the rotor flowfield are computed including the flow acceleration through the rotor disk, the axial and swirl velocities in the rotor downwash, and the slipstream contraction. Many of the complex tilt rotor flow features are captured including the highly three-dimensional flow over the wing, the recirculation fountain at the plane of symmetry, wing leading and trailing edge separation, and the large region of separated flow beneath the wing. Mean wing surface pressures compare fairly well with available experimental data, but the time-averaged download/thrust ratio is 20-30 percent higher than the measured value. The discrepancy is due to a combination of factors that are discussed. Leading edge tangential blowing, of constant strength along the wing span, is shown to be effective in reducing download. The jet serves primarily to reduce the pressure on the wing upper surface. The computation clearly shows that, because of the three-dimensionality of the flowfield, optimum blowing would involve a spanwise variation in blowing strength.

Fejtek, Ian G.↗
NASA NTRS1959

Experimental Investigation of a 0.35 Hub-Tip Radius Ratio Transonic Axial Flow Rotor Designed for 40 Pounds per Second per Square Foot with a Design Tip Diffusion Factor of 0.20

In order to determine the effect of a low design diffusion factor on the performance of a transonic axial-flow compressor rotor, a high-specific-flow rotor with a 0.35 hub-tip radius ratio was designed, fabricated and tested. This rotor used a design tip diffusion factor of 0.20 with a design corrected specific weight flow of 40 pounds per second per square foot of frontal area, a total-pressure ratio of 1.27, and an adiabatic efficiency of 0.96. The design, rotor performance, and blade element performance are presented with a discussion on rotor shock losses and a comparison with a similarly designed rotor with a tip diffusion factor of 0.35. At the design corrected tip speed of 1100 feet per second, a peak rotor adiabatic efficiency of 0.88 was attained at a corrected specific weight flow of 39 pounds per second per square foot of frontal area with a mass-averaged total-pressure ratio of 1.27. The blade element tip diffusion factor was 0.281, which is 0.08 higher than the design value of 0.20. Peak efficiencies of 0.95, 0.91, 0.89, and 0.85 were obtained at 70, 80, 90, and 110 percent of design speed, respectively. Comparison of the performance of the rotor reported herein and a similarly designed rotor with increased blade loading indicates that higher blade loading results in a more desirable rotor because of a higher pressure ratio and equivalent efficiency. Computed values of shock losses at the rotor tip section indicate that the losses at peak efficiency are primarily a function of shock losses since the profile losses are only a small percentage of the total loss.

Yasaki, Paul T.↗
NASA NTRS1960

Hovering Measurements for Twin-Rotor Configurations with and Without Overlap

Results of an investigation in the Langley full-scale tunnel of the hovering performance of large-scale twin-rotor-helicopter models are presented. Measurements of thrust, torque, and rotor flapping are given for overlapped (approximately 76 percent of blade radius) and nonoverlapped configurations and for two different rotor solidities. The measured performance is compared with single-rotor measurements and with available rotor theory. These tests show that the hovering performance of a single rotor and of two rotors without overlap or vertical offset are the same and hence may be calculated by single-rotor theory. These tests in conjunction with results of previous coaxial-rotor tests show that the performance of highly overlapped rotors can be reasonably predicted by available rotor theory.

Sweet, George E.↗
NASA NTRS1961

Some Recent Studies in Structural Dynamics of Rotor Aircraft

This paper presents a summary of four recent studies relating to the structural-dynamics problems of rotor-powered aircraft. The first study concerns the measurement by means of dynamic models of the forces and moments at the hubs of various rotor configurations. The results of this study show that the periodic components of the forces and moments are highly dependent on both the rotor configuration and the flight condition. The second study treats the problem of resonance amplifications of rotor-blade stress and shows that by using multiple flapping hinges or flex-joints it is possible to control the natural frequencies of the rotor blade so that conditions of resonance between the frequencies of the aerodynamic input forces and the natural frequencies of the lower blade modes are avoided for all rotor speeds. Two studies of the stability of rotor aircraft axe also discussed. One of these involves the mechanical instability or ground resonance of rotorcraft wherein the rotor support in each of two mutually perpendicular directions in the rotor plane is represented as a multiple-degree-of-freedom system in contrast to the system having a "single" degree of freedom normally used in helicopter analysis. The consideration of the rotor support system as a two-degree-of-freedom system predicts additional unstable ranges of rotor speed not predicted by former analyses. The other instability treated is propeller whirl for which the significant motions are the pitching and yawing motions of the propeller disk which are coupled together by gyroscopic forces. The effect of the significant variables on the instability is presented and shows that the speed at which the instability occurs can be increased by increasing the relative stiffness of the propeller in the pitch and yaw directions or by increasing the effective structural damping of the nacelle.

George W Brooks↗
NASA NTRS2015

Wind Tunnel Interference Effects on Tilt Rotor Testing Using Computational Fluid Dynamics

Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tilt Rotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity URANS solver is used with an incompressible flow model and a realizable k-ε turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade element model with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at NASA Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A 'quasi linear trim' was used to trim the thrust for the rotor to compare the power as a unique variable. Power differences between free field and wind tunnel cases were found from -7 % to 0 % in the 80- by 120-Foot Wind Tunnel test section and -1.6 % to 4.8 % in the 40- by 80-Foot Wind Tunnel, depending on the TTR orientation, tunnel velocity and blade setting. The TTR will be used in 2016 to test the Bell 609 rotor in a similar fashion to the research in this report.

CFD↗