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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Effect of Variable Chord Length on Transonic Axial Rotor Performance Investigated

During the life of any gas turbine, blade erosion is present, especially for those units that are exposed to unfiltered air, such as aviation turbofan engines. The effect of this erosion is to reduce the blade chord progressively from the midspan to the tip region and to roughen and distort the blade surface. The effects of roughness on rotor performance have been documented by Suder et al. and Roberts. These papers indicate that the penalty for leading-edge roughness and erosion can be significant. Turbofan operators, therefore, restore chord length at routine maintenance intervals to regain performance before deterioration is too severe to salvage blades. As the rotor blades erode, the leading edge becomes rough - blunt and distorted from the nominal shape - and the aerodynamic performance suffers. Nominal performance can be recovered by recontouring the leading edges. This process, which inherently shortens the blade chord, can be used until the blade chord erodes to the stall limit. Below this chord length, which varies among engine-compressor types, a decrease of stall margin is likely. After compressor blade rework that includes leading edge recontouring, the blades have different chord lengths, ranging from blades that are near nominal chord length down to those near the stall chord limit. Furthermore, as blades erode below the stall limit, they must be replaced with new blades that have the full nominal chord length. Consequently, a set of compressor blades with varying chord lengths will be installed into each turbofan engine that goes through a complete maintenance cycle. The question arises, "Does fan or compressor performance depend on the order in which mixed-chord blades are installed into a fan or compressor disk?"

Suder, Kenneth L.↗
NASA NTRS2002

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

Roberts, William B.↗
NASA NTRS1990

Exact Chord-Length Distribution For SEU Calculations

Computed rates of SEU's more accurate. Exact integral chord-length distribution derived for use in calculations of rates of single-event upsets (SEU's) (changes in logic states) caused by impingement of cosmic rays or other ionizing radiation on electronic logic circuits.

Buehler, Martin G.↗
NASA NTRS1988

An exact, closed-form expression of the integral chord-length distribution for the calculation of single-event upsets induced by cosmic rays

This paper presents a derivation of an exact closed-form expression of the integral chord-length distribution for the calculation of single-event upsets (SEUs) in an electronic memory cell, caused by cosmic rays. Results computed for two rectangular parallelepipeds using this exact expression are compared with those computed with Bradford's (1979) semiexact expression of C(x). It is found that the values of C(x) are identical for x equal or smaller than b but are vastly different for x greater than b. Moreover, while C(x) of Bradford gives reasonably accurate values of SEU rate for certain sets of computational parameters, it gives values more than 10 times larger than the correct values for other sets of parameters.

Luke, Keung L.↗
NASA NTRS1991

Development of a perturbation generator for vortex stability studies

Theory predicts vortex instability when subjected to certain types of disturbances. It was desired to build a device which could introduce controlled velocity perturbations into a trailing line vortex in order to study the effects on stability. A perturbation generator was designed and manufactured which can be attached to the centerbody of an airfoil type vortex generator. Details of design tests and manufacturing of the perturbation generator are presented. The device produced controlled perturbation with frequencies in excess of 250 Hz. Preliminary testing and evaluation of the perturbation generator performance was conducted in a 4 inch cylindrical pipe. Observations of vortex shedding frequencies from a centerbody were measured. Further evaluation with the perturbation generator attached to the vortex generator in a 2 x 3 foot wind tunnel were also conducted. Hot-wire anemometry was used to confirm the perturbation generator's ability to introduce controlled frequency fluctuations. Comparison of the energy levels of the disturbances in the vortex core was made between locations 42 chord lengths and 15 chord lengths downstream.

Riester, J. E.↗
NASA NTRS2003

Design-of-experiments to Reduce Life-cycle Costs in Combat Aircraft Inlets

It is the purpose of this study to demonstrate the viability and economy of Design- of-Experiments (DOE), to arrive at micro-secondary flow control installation designs that achieve optimal inlet performance for different mission strategies. These statistical design concepts were used to investigate the properties of "low unit strength" micro-effector installation. "Low unit strength" micro-effectors are micro-vanes, set a very low angle-of incidence, with very long chord lengths. They are designed to influence the neat wall inlet flow over an extended streamwise distance. In this study, however, the long chord lengths were replicated by a series of short chord length effectors arranged in series over multiple bands of effectors. In order to properly evaluate the performance differences between the single band extended chord length installation designs and the segmented multiband short chord length designs, both sets of installations must be optimal. Critical to achieving optimal micro-secondary flow control installation designs is the understanding of the factor interactions that occur between the multiple bands of micro-scale vane effectors. These factor interactions are best understood and brought together in an optimal manner through a structured DOE process, or more specifically Response Surface Methods (RSM).

Anderson, Bernhard H.↗
NASA NTRS2009

Subspace Iteration Method for Complex Eigenvalue Problems with Nonsymmetric Matrices in Aeroelastic System

Modern airplane design is a multidisciplinary task which combines several disciplines such as structures, aerodynamics, flight controls, and sometimes heat transfer. Historically, analytical and experimental investigations concerning the interaction of the elastic airframe with aerodynamic and in retia loads have been conducted during the design phase to determine the existence of aeroelastic instabilities, so called flutter .With the advent and increased usage of flight control systems, there is also a likelihood of instabilities caused by the interaction of the flight control system and the aeroelastic response of the airplane, known as aeroservoelastic instabilities. An in -house code MPASES (Ref. 1), modified from PASES (Ref. 2), is a general purpose digital computer program for the analysis of the closed-loop stability problem. This program used subroutines given in the International Mathematical and Statistical Library (IMSL) (Ref. 3) to compute all of the real and/or complex conjugate pairs of eigenvalues of the Hessenberg matrix. For high fidelity configuration, these aeroelastic system matrices are large and compute all eigenvalues will be time consuming. A subspace iteration method (Ref. 4) for complex eigenvalues problems with nonsymmetric matrices has been formulated and incorporated into the modified program for aeroservoelastic stability (MPASES code). Subspace iteration method only solve for the lowest p eigenvalues and corresponding eigenvectors for aeroelastic and aeroservoelastic analysis. In general, the selection of p is ranging from 10 for wing flutter analysis to 50 for an entire aircraft flutter analysis. The application of this newly incorporated code is an experiment known as the Aerostructures Test Wing (ATW) which was designed by the National Aeronautic and Space Administration (NASA) Dryden Flight Research Center, Edwards, California to research aeroelastic instabilities. Specifically, this experiment was used to study an instability known as flutter. ATW was a small-scale airplane wing comprised of an airfoil and wing tip boom. This wing was formulated based on a NACA-65A004 airfoil shape with a 3.28 aspect ratio. The wing had a span of 18 inch with root chord length of 13.2 inch and tip chord length of 8.7 inch. The total area of this wing was 197 square inch. The wing tip boom was a 1 inch diameter hollow tube of length 21.5 inch. The total weight of the wing was 2.66 lbs.

Pak, Chan-gi↗
NASA NTRS1991

Lift producing device exhibiting low drag and reduced ventilation potential and method for producing the same

A lift producing device is disclosed which is adapted to be connected to a vehicle to provide lift to the vehicle when the vehicle is moved relative to a first fluid medium having a first density and viscosity and being in contact with a second fluid medium adjacent the vehicle. The second fluid medium has a second fluid density which is different from the first fluid density. The lift producing device comprises opposed first and second major surfaces joined at a longitudinally extending leading edge and at a longitudinally extending trailing edge, with at least a portion of the longitudinally extending leading edge being spaced from the longitudinally extending trailing edge by a predetermined mean chord length. When the vehicle is moved relative to the first fluid medium at a velocity within a range of predetermined velocities, with each of the velocities having a direction inclined from a plane extending through the leading edge and the trailing edge within a predetermined angular range, a region of high pressure is generated in the first fluid medium adjacent the first major surface and a region of low pressure is generated in the first fluid medium adjacent the second major surface. The lift producing device has a cross-sectional shape which will generate a pressure distribution around the device when the vehicle is moved relative to the first fluid medium at a velocity within the range of predetermined velocities such that the first fluid medium exhibits attached laminar flow along the device for a portion of the predetermined mean chord length from the leading edge to the trailing edge and will neither form a laminar separation bubble adjacent the second major surface of the device, nor exhibit turbulent separation adjacent the second major surface for substantially all of the predetermined mean chord length from the leading edge to the trailing edge. The portion along which attached laminar flow is maintained is the longest portion which will still fulfill the flow separation requirements. A method for producing the foil is also disclosed.

Caldwell, Richard A.↗
NASA NTRS1989

Compression Pylon

A compression pylon for an aircraft with a wing-mounted engine, that does not cause supersonic airflow to occur within the fuselage-wing-pylon-nacelle channel is presented. The chord length of the pylon is greater than the local chord length of the wing to which it is attached. The maximum thickness of the pylon occurs at a point corresponding to the local trailing edge of the wing. As a result, the airflow through the channel never reaches supersonic velocities.

Patterson, James C., Jr.↗
NASA NTRS1999

Spanwise transition section for blended wing-body aircraft

A blended wing-body aircraft includes a central body, a wing, and a transition section which interconnects the body and the wing on each side of the aircraft. The two transition sections are identical, and each has a variable chord length and thickness which varies in proportion to the chord length. This enables the transition section to connect the thin wing to the thicker body. Each transition section has a negative sweep angle.

Hawley, Arthur V.↗
NASA NTRS2004

Assessment of Comprehensive Analysis Calculation of Airloads on Helicopter Rotors

Blade section normal force and pitching moment were investigated for six rotors operating at transition and high speeds: H-34 in flight and wind tunnel, SA 330 (research Puma), SA 349/2, UH-60A full-scale and BO-105 model (HART-I). The measured data from flight and wind tunnel tests were compared with calculations obtained using the comprehensive analysis CAMRAD II. The calculations were made using two free wake models: rolled-up and multiple-trailer with consolidation models. At transition speed, there is fair to good agreement for the blade section normal force between the test data and analysis for the H-34, research Puma, and SA 349/2 with the rolled-up wake. The calculated airloads differ significantly from the measurements for the UH-60A and BO-105. Better correlation is obtained for the UH-60A and BO-105 by using the multiple-trailer with consolidation wake model. In the high speed condition, the analysis shows generally good agreement with the research Puma flight data in both magnitude and phase. However, poor agreement is obtained for the other rotors examined. The analysis shows that the aerodynamic tip design (chord length and quarter chord location) of the Puma has an important influence on the phase correlation.

Yeo, Hyeonsoo↗
NASA NTRS2026

Aerodynamic Performance and Acoustic Impacts of Varying Tip Speeds and Tripping Conditions on Small Rotors in an Anechoic Hover Chamber

Performance and acoustic measurements were taken in a hover chamber for various optimum hovering rotors (OPT2) and a commercial-off-the-shelf (COTS) rotor. A total of 10 rotors are included in this report, all of which have two blades and a tip radius of 0.1905 m (7.5 in.). For the OPT2, results for three additive manufacturing methods of fabrication are presented: stereolithography(SLA) using Accura Xtreme, SLA using FormLabs 10K resin, and selective laser sintering (SLS) using mineral-filled PA12 nylon material. All but one set of rotors were designed with a trailing edge bluntness that is 3% of the chord length, and one set was designed with a bluntness that is 1% of the chord length. Spanwise boundary layer trips were applied to the SLA, FormLabs and COTS rotors. Performance comparisons between untripped and tripped configurations demonstrate the impact of boundary layer state on rotor efficiency. Acoustic results, including periodic and broadband noise components, are presented. The effects of tripping these rotors near the leading edge are also presented. For some of the rotors, acoustic spectra of tip speed sweeps are presented to show how the frequency content changes with tip speed. Amplitude and frequency scaling methods are used to collapse broadband spectra at various tip speeds towards a unified curve. These findings contribute to the understanding of small rotor aeroacoustics and provide valuable datasets for computational model validation in urban air mobility applications.

eVTOL↗
NASA NTRS1960

A Wind-Tunnel Investigation of the Development of Lift on Wings in Accelerated Longitudinal Motion

An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the development of lift on a wing during a simulated constant-acceleration catapult take-off. The investigation included models of a two-dimensional wing, an unswept wing having an aspect ratio of 6, a 35 deg. swept wing having an aspect ratio of 3.05, and a 60 deg. delta wing having an aspect ratio of 2.31. All the wings investigated developed at least 90 percent of their steady-state lift in the first 7 chord lengths of travel. The development of lift was essentially independent of the acceleration when based on chord lengths traveled, and was in qualitative agreement with theory.

Turner, Thomas R.↗