SEARCH · Search NASA

Results for “ROTOR DISK”

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.

SourceResearch connectionFrom yearThrough year

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

At least 55 records · Page 3

All sources NASA NTRS DOE OSTI

NASA NTRS1977

Steady-state unbalance response of a three-disk flexible rotor on flexible, damped supports

Experimental data are presented for the unbalance response of a flexible, ball bearing supported rotor to speeds above the third lateral bending critical. Values of squeeze film damping coefficients obtained from measured data are compared to theoretical values obtained from short bearing approximation over a frequency range from 5000 to 31,000 cycles/min. Experimental response for an undamped rotor is compared to that of one having oil squeeze film dampers at the bearings. Unbalances applied varied from 0.62 to 15.1 gm-cm.

Cunningham, R. E.↗

NASA NTRS1997

Rotor-Fuselage Interaction: Analysis and Validation with Experiment

The problem of rotor-fuselage aerodynamic interaction has to be considered in industry applications from various aspects. First, in order to increase helicopter speed and reduce operational costs, rotorcraft tend to be more and more compact, with a main rotor closer to the fuselage surface. This creates significant perturbations both on the main rotor and on the fuselage, including steady and unsteady effects due to blade and wake passage and perturbed inflow at the rotor disk. Furthermore,the main rotor wake affects the tail boom, empennage and anti-torque system. This has important consequences for helicopter control and vibrations at low speeds and also on tail rotor acoustics (main rotor wake-tail rotor interactions). This report describes the US Army-France MOD cooperative work on this problem from both the theoretical and experimental aspects. Using experimental 3D velocity field and fuselage surface pressure measurements, three codes that model the interactions of a helicopter rotor with a fuselage are compared. These comparisons demonstrate some of the strengths and weaknesses of current models for the combined rotor-fuselage analysis.

Berry, John D.↗

NASA NTRS2010

Conceptual Design of a Two Spool Compressor for the NASA Large Civil Tilt Rotor Engine

This paper focuses on the conceptual design of a two spool compressor for the NASA Large Civil Tilt Rotor engine, which has a design-point pressure ratio goal of 30:1 and an inlet weight flow of 30.0 lbm/sec. The compressor notional design requirements of pressure ratio and low-pressure compressor (LPC) and high pressure ratio compressor (HPC) work split were based on a previous engine system study to meet the mission requirements of the NASA Subsonic Rotary Wing Projects Large Civil Tilt Rotor vehicle concept. Three mean line compressor design and flow analysis codes were utilized for the conceptual design of a two-spool compressor configuration. This study assesses the technical challenges of design for various compressor configuration options to meet the given engine cycle results. In the process of sizing, the technical challenges of the compressor became apparent as the aerodynamics were taken into consideration. Mechanical constraints were considered in the study such as maximum rotor tip speeds and conceptual sizing of rotor disks and shafts. The rotor clearance-to-span ratio in the last stage of the LPC is 1.5% and in the last stage of the HPC is 2.8%. Four different configurations to meet the HPC requirements were studied, ranging from a single stage centrifugal, two axi-centrifugals, and all axial stages. Challenges of the HPC design include the high temperature (1,560deg R) at the exit which could limit the maximum allowable peripheral tip speed for centrifugals, and is dependent on material selection. The mean line design also resulted in the definition of the flow path geometry of the axial and centrifugal compressor stages, rotor and stator vane angles, velocity components, and flow conditions at the leading and trailing edges of each blade row at the hub, mean and tip. A mean line compressor analysis code was used to estimate the compressor performance maps at off-design speeds and to determine the required variable geometry reset schedules of the inlet guide vane and variable stators that would result in the transonic stages being aerodynamically matched with high efficiency and acceptable stall margins based on user specified maximum levels of rotor diffusion factor and relative velocity ratio.

Veres, Joseph P.↗

NASA NTRS2004

Blade-Pitch Control for Quieting Tilt-Rotor Aircraft

A method of reducing the noise generated by a tilt-rotor aircraft during descent involves active control of the blade pitch of the rotors. This method is related to prior such noise-reduction methods, of a type denoted generally as higher-harmonic control (HHC), in which the blade pitch is made to oscillate at a harmonic of the frequency of rotation of the rotor. A tilt-rotor aircraft is so named because mounted at its wing tips are motors that can be pivoted to enable the aircraft to take off and land like a helicopter or to fly like a propeller airplane. When the aircraft is operating in its helicopter mode, the rotors generate more thrust per unit rotor-disk area than helicopter rotors do, thus producing more blade-vortex interaction (BVI) noise. BVI is a major source of noise produced by helicopters and tilt-rotor aircraft during descent: When a rotor descends into its own wake, the interaction of each blade with the blade-tip vortices generated previously gives rise to large air-pressure fluctuations. These pressure fluctuations radiate as distinct, impulsive noise. In general, the pitch angle of the rotor blades of a tilt-rotor aircraft is controlled by use of a swash plate connected to the rotor blades by pitch links. In both prior HHC methods and the present method, HHC control signals are fed as input to swash-plate control actuators, causing the rotor-blade pitch to oscillate. The amplitude, frequency, and phase of the control signal can be chosen to minimize BVI noise.

Betzina, Mark D.↗

NASA NTRS2002

Rotor Wake Vortex Definition: Initial Evaluation of 3-C PIV Results of the Hart-II Study

An initial evaluation is made of extensive three-component (3C) particle image velocimetry (PIV) measurements within the wake across a rotor disk plane. The model is a 40 percent scale BO-105 helicopter main rotor in forward flight simulation. This study is part of the HART II test program conducted in the German-Dutch Wind Tunnel (DNW). Included are wake vortex field measurements over the advancing and retreating sides of the rotor operating at a typical descent landing condition important for impulsive blade-vortex interaction (BVI) noise. Also included are advancing side results for rotor angle variations from climb to steep descent. Using detailed PIV vector maps of the vortex fields, methods of extracting key vortex parameters are examined and a new method was developed and evaluated. An objective processing method, involving a center-of-vorticity criterion and a vorticity 'disk' integration, was used to determine vortex core size, strength, core velocity distribution characteristics, and unsteadiness. These parameters are mapped over the rotor disk and offer unique physical insight for these parameters of importance for rotor noise and vibration prediction.

Burley, Casey L.↗

NASA NTRS1976

Parametric study of the noise produced by the interaction of the main rotor wake with the tail rotor

A model was designed, fabricated and wind tunnel tested to identify some of the parameters which were pertinent to the noise produced by the interaction of the main rotor wake with the tail rotor. The model provided for variations in many geometric and operating parameters. The initial set of tests indicated that the noise produced by the tail rotor was, in general, sensitive to the location of the vortex interaction on the tail rotor disk, direction of rotation, lateral rotor fin spacing, tip speed and the operating mode of the tail rotor; and generally insensitive to main rotor thrust coefficient, longitudinal spacing and tail rotor to main rotor rotational speed ratios. Refinements in the analyses to adequately predict the noise phenomenon have been outlined to complement further experimental investigations.

Balcerak, J. C.↗

NASA NTRS2019

Overflow Analysis of Unmanned Aircraft Systems

With the increasing presence of unmanned aircraft systems (UAS), or drones, in the national airspace, the management for access and operation of these vehicles is required. NASA’s management approach is being developed under the unmanned aircraft system traffic management (UTM) program. To determine the aerodynamic characteristics of drones, wind tunnel experiments and computation fluid dynamic (CFD) analysis have been conducted. These experiments and analyses are undertaken to understand the flight capabilities of these vehicles in variable head and cross wind conditions. The results of these investigations will provide metrics for the safe operation of these vehicles in and around civil populations and in urban settings. The focus of this paper is to model a drone installed in a wind tunnel for varying pitch attitudes and rotor rpm settings. Specifically, the IRIS drone is modeled in the NASA-Ames 7x10 ft wind tunnel. The tunnel mounting hardware and the tunnel enclosure are modeled along with the IRIS drone geometry. The rotors of the drone are modeled using two methodologies: a rotor disk model and full rotating rotors with moving grids. The results of the analysis are compared with available experimental data to validate the computational approach.

Stremel, Paul M.↗

NASA NTRSDate not supplied

Multi-Fidelity Computational Analysis of a Quiet Single-Main Rotor Helicopter for Air Taxi Operations

A multi-fidelity computational study has been conducted on NASA’s quiet single-main rotor helicopter concept for urban air mobility using computational fluid dynamics and rotorcraft comprehensive analysis tools. Two approaches have been followed in order to simulate the rotor flows: (1) a rotor disk model, which models the rotors as source terms, and (2) a high-fidelity approach using high-order accurate schemes and dual-time stepping, which simulates the rotor with its individual rotating blade grids. The delayed detached-eddy simulation model has been employed. A loose-coupling approach between the flow solver and a rotorcraft comprehensive code is utilized to include vehicle trim and motion variables. The quiet single-main rotor helicopter is one of the conceptual designs intended to focus and guide NASA’s research activities in support of aircraft development for vertical take-off and landing air taxi operations.

RVLT↗

NASA NTRS1985

Aerodynamic interaction between vortical wakes and the viscous flow about a circular cylinder

In the design analysis of conventional aircraft configurations, the prediction of the strong interaction between vortical wakes and the viscous flow field about bodies is of considerable importance. Interactions between vortical wakes and aircraft components are even more common on rotorcraft and configurations with lifting surfaces forward of the wing. An accurate analysis of the vortex-wake interaction with aircraft components is needed for the optimization of the payload and the reduction of vibratory loads. However, the three-dimensional flow field beneath the rotor disk and the interaction of the rotor wake with solid bodies in the flow field are highly complex. The present paper has the objective to provide a basis for the considered interactions by studying a simpler problem. This problem involves the two-dimensional interaction of external wakes with the viscous flow about a circular cylinder.

Stremel, P. M.↗

NASA NTRS1985

Thrust changes induced by ground and ceiling planes on a rotor in hover

The thrust produced by a helicopter motor in hover has been studied by measuring the thrust on a model rotor 0.234 meters in diameter operating between ground and ceiling planes. The distance between the two planes and the rotor disk was in the range 0.08-6.0 rotor diameters. It was found that the confining surfaces changed the rotor thrust almost linearly with the logarithm of the distance from the confining surface. When both planes were near the rotor the variation was no longer linear. The measured thrust of the rotor was compared to predictions based on the theoretical calculations of Rossow (1985), and the results are given in a table.

Rossow, V. J.↗

DOE OSTI2023

Large Eddy Simulation of Rotationally Induced Ingress and Egress around an Axial Seal between Rotor and Stator Disks

In gas turbines, the hot gas exiting the combustor can have temperatures as high as 2000 °C, and some of this hot gas enter into the space between the stator and rotor disks (wheelspace). Since the entering hot gas could damage the disks, its ingestion must be minimized. This is carried out by rim seals and by introducing a cooler flow from the compressor (sealing flow) into the wheelspace. Ingress and egress into rim seals are driven by the stator vanes, the rotor and its rotation, and the rotor blades. This study focuses on the ingress and egress driven by the rotor and its rotation. This is carried out by performing wall-resolved large eddy simulation (LES) around an axial seal in a rotor–stator configuration without vanes and blades. Results obtained show the mechanisms by which the rotor and its rotation induce ingress, egress, and flow trajectories. Kelvin–Helmholtz instability was found to create a wavy shear layer and displacement thickness that produces alternating regions of high and low pressures around the rotor side of the seal. Vortex shedding on the backward-facing side of the seal and its impingement on the rotor side of the seal also produces alternating regions of high and low pressures. The locations of the alternating regions of high and low pressures were found to be statistically stationary and to cause ingress to start on the rotor side of the seal. Vortex shedding and recirculating flow in the seal clearance also cause ingress by entrainment. With the effects of the rotor and its rotation on ingress and egress isolated, this study enables the effects of stator vanes and rotor blades to be assessed.

42 ENGINEERING↗

NASA NTRS1976

A parametric model study of the noise generated by the aerodynamic interaction of the tail rotor with the wake of the main rotor

Wind tunnel tests were conducted to determine the influence of various parameters on the noise produced by the interaction of the main rotor wake with the tail rotor. Parameters found to be of significance were the location of the main rotor vortex interaction on the tail rotor disk, the direction of rotation of the tail rotor, the lateral tail rotor-fin spacing, and the operating mode of the tail rotor. Of lesser importance were the main rotor thrust coefficient, longitudinal spacing between rotors, and tail rotor-to-main rotor rotational speed ratios. Analysis of the interaction noise characteristics using a simplified representation of the aerodynamic phenomena occurring during the interaction showed that existing predictive techniques could be modified and extended to predict the pertinent noise characteristics.

White, R. P., Jr.↗

NASA NTRS2015

Experimental Investigation of Rotorcraft Outwash in Ground Effect

The wake characteristics of a rotorcraft are affected by the proximity of a rotor to the ground surface, especially during hover. Ground effect is encountered when the rotor disk is within a distance of a few rotor radii above the ground surface and results in an increase in thrust for a given power relative to that same power condition with the rotor out of ground effect. Although this phenomenon has been highly documented and observed since the beginning of the helicopter age, there is still a relatively little amount of flow-field data existing to help understand its features. Joint Army and NASA testing was conducted at NASA Langley Research Center using a powered rotorcraft model in hover at various rotor heights and thrust conditions in order to contribute to the complete outwash data set. The measured data included outwash velocities and directions, rotor loads, fuselage loads, and ground pressures. The researchers observed a linear relationship between rotor height and percent download on the fuselage, peak mean outwash velocities occurring at radial stations between 1.7 and 1.8 r/R regardless of rotor height, and the measurement azimuthal dependence of the outwash profile for a model incorporating a fuselage. Comparisons to phase-locked PIV data showed similar contours but a more contracted wake boundary for the PIV data. This paper describes the test setup and presents some of the averaged results.

Tanner, Philip E.↗

NASA NTRS2022

Identification and Reduction of Interactional Noise of a Quadcopter in Hover and Forward Flight Conditions

Advanced Air Mobility is a vision for a safe, accessible, and sustainable aviation system to transport people and cargo between places not served by traditional aviation. With this emerging transportation industry, there is motivation to characterize the noise of vehicles to determine their potential impacts on the community. An experimental testing campaign was conducted on a representative model of a small unmanned aircraft system in the NASA Langley Low Speed Aeroacoustic Wind Tunnel as a continuation of a previous testing campaign. The goals of the current test are to identify sources of interactional noise as well as to test custom-designed rotors and noise reduction devices. The tested noise reduction methods involve increasing the vertical distances between the rotors and the vehicle airframe as well as between the forward and aft rotor disk planes. These methods are intended to reduce rotor-airframe interaction noise in hover and fore-aft rotor wake ingestion noise in forward flight. A phased microphone array is also utilized to identify the locations of prominent noise generation for the different vehicle configurations in forward flight. Elevation of the rotors from the vehicle airframe yielded nearly 8 dBA overall noise reduction in forward flight, while yielding up to 4 dB reduction in overall tonal levels for one of the rotors in hover.

Nikolas S Zawodny↗

NASA NTRS2008

Influence of Lift Offset on Rotorcraft Performance

The influence of lift offset on the performance of several rotorcraft configurations is explored. A lift-offset rotor, or advancing blade concept, is a hingeless rotor that can attain good efficiency at high speed, by operating with more lift on the advancing side than on the retreating side of the rotor disk. The calculated performance capability of modern-technology coaxial rotors utilizing a lift offset is examined, including rotor performance optimized for hover and high-speed cruise. The ideal induced power loss of coaxial rotors in hover and twin rotors in forward flight is presented. The aerodynamic modeling requirements for performance calculations are evaluated, including wake and drag models for the high speed flight condition. The influence of configuration on the performance of rotorcraft with lift-offset rotors is explored, considering tandem and side-by-side rotorcraft as well as wing-rotor lift share.

Johnson, Wayne↗

NASA NTRS2009

Influence of Lift Offset on Rotorcraft Performance

The influence of lift offset on the performance of several rotorcraft configurations is explored. A lift-offset rotor, or advancing blade concept, is a hingeless rotor that can attain good efficiency at high speed by operating with more lift on the advancing side than on the retreating side of the rotor disk. The calculated performance capability of modern-technology coaxial rotors utilizing a lift offset is examined, including rotor performance optimized for hover and high-speed cruise. The ideal induced power loss of coaxial rotors in hover and twin rotors in forward flight is presented. The aerodynamic modeling requirements for performance calculations are evaluated, including wake and drag models for the high-speed flight condition. The influence of configuration on the performance of rotorcraft with lift-offset rotors is explored, considering tandem and side-by-side rotorcraft as well as wing-rotor lift share.

Johnson, Wayne↗

DOE OSTI2024

Investigation of a Multi-Rotor Triboelectric Nanogenerator Using a Modular Flexible Circuit Board Stack

Triboelectric nanogenerators (TENGs) are a nascent class of energy harvester that are being explored for scavenging energy from small ocean waves. To date, they have been integrated in wave energy converters (WECs) designed to capture random motion caused by the perturbations of the ocean surface. Blue economy applications such as ocean observation can benefit greatly from more substantial wave-derived power, as such, the power output of existing TENG WECs must be increased several-fold to become viable. This study describes the conceptualization of a rotary TENG, and the subsequent efforts to increase its power output by stacking multiple stator and rotor pairs. In the latter part of this study, we introduce a novel means of incorporating a friction element into the design of the TENG by employing flexible printed circuit board (PCB) rotors. At rest, these flexible rotors will contact the stator, building static charges due to friction, at speed, these flexible rotors will be decoupled from the stator, reducing friction and allowing faster rotation. Initial results indicate that the power output of the flexible PCB rotor does not produce more power than a rigid, acrylic disk rotor, however the current output of the prototype is boosted, and the prototype is far more compact, allowing for a higher energy density than the rigid rotor prototype.

bench testing↗

NASA NTRS1990

Probabilistic analysis of bladed turbine disks and the effect of mistuning

Probabilistic assessment of the maximum blade response on a mistuned rotor disk is performed using the computer code NESSUS. The uncertainties in natural frequency, excitation frequency, amplitude of excitation and damping are included to obtain the cumulative distribution function (CDF) of blade responses. Advanced mean value first order analysis is used to compute CDF. The sensitivities of different random variables are identified. Effect of the number of blades on a rotor on mistuning is evaluated. It is shown that the uncertainties associated with the forcing function parameters have significant effect on the response distribution of the bladed rotor.

Shah, A. R.↗