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NASA NTRS1972

An investigation of compressor noise by use of a spinning mode synthesizer.

Use of the Spinning Mode Synthesizer (SMS) as a tool for studying the properties of such complex waves as those generated by fan engines. This has greatly reduced the amount of trial and error techniques used for optimizing duct liners. The SMS was designed by using simple transducer theory and properly controlling the phasing to each of 8 loudspeakers. It provides a means for making in-duct measurements of the radial, axial, and circumferential modal distributions. Results are presented which show that the device is of value for investigation of the highly complex nature of compressor radiated noise.

Oslac, M. J.↗

NASA NTRS2014

Arbitrary Shape Deformation in CFD Design

Sculptor(R) is a commercially available software tool, based on an Arbitrary Shape Design (ASD), which allows the user to perform shape optimization for computational fluid dynamics (CFD) design. The developed software tool provides important advances in the state-of-the-art of automatic CFD shape deformations and optimization software. CFD is an analysis tool that is used by engineering designers to help gain a greater understanding of the fluid flow phenomena involved in the components being designed. The next step in the engineering design process is to then modify, the design to improve the components' performance. This step has traditionally been performed manually via trial and error. Two major problems that have, in the past, hindered the development of an automated CFD shape optimization are (1) inadequate shape parameterization algorithms, and (2) inadequate algorithms for CFD grid modification. The ASD that has been developed as part of the Sculptor(R) software tool is a major advancement in solving these two issues. First, the ASD allows the CFD designer to freely create his own shape parameters, thereby eliminating the restriction of only being able to use the CAD model parameters. Then, the software performs a smooth volumetric deformation, which eliminates the extremely costly process of having to remesh the grid for every shape change (which is how this process had previously been achieved). Sculptor(R) can be used to optimize shapes for aerodynamic and structural design of spacecraft, aircraft, watercraft, ducts, and other objects that affect and are affected by flows of fluids and heat. Sculptor(R) makes it possible to perform, in real time, a design change that would manually take hours or days if remeshing were needed.

Landon, Mark↗

NASA NTRS2008

Fan Acoustic Issues in the NASA Space Flight Experience

Emphasis needs to be placed on choosing quiet fans compatible with systems design and specifications that control spec levels: a) Sound power; b) Choose quiet fan or plan to quiet it, early in program; c) Plan early verification that fan source allocations are met. Airborne noise: a) System design should function/play together with fans used (flow passages, restrictions, bends, expansions & contractions, and acoustics) vs. fan speed understood (nominal, worst case, & unplanned variances); b) Fan inlets treated, as required; c) Fan Outlets treated, as required; d) Ducted system inlets are outlets designed for acoustic compliance compatibility & designed so some late required modifications can be made without significant impacts. Structure Borne Noise: a) Structure borne noise dealt with as part of fan package or installation; b) Duct attachments and lines isolated. Case Radiated Noise: - Treatment added as much as possible to fan package (see example).

Allen, Christopher S.↗

NASA NTRS1999

Comparison of CFD Predictions of the TCA Baseline

The computational fluid dynamics (CFD) comparisons being presented are compared to each other and to wind tunnel (WT) data on the baseline TCA. Some of the CFD computations were done prior to the tests and others later. Only force data (CL vs CD) from CFD will be presented as part of this report. The WT data presented comes from the testing of the baseline TCA in the Langley Unitary Plan Wind Tunnel (UPWT), Test Section #2. There are 2 sets of wind tunnel data being presented: one from test 1671 of model 2a (flapped wing) and the other from test 1679 of model 2b (solid wing). Most of the plots show only one run from each of the WT tests per configuration. But many repeat runs were taken during the tests. The WT repeat runs showed an uncertainty in the drag of +/- 0.5 count. There were times when the uncertainty in drag was better, +/- 0.25 count. Test 1671 data was of forces and pressures measured from model 2a. The wing had cutouts for installing various leading and trailing edge flaps at lower Mach numbers. The internal duct of the nacelles are not designed and fabricated as defined in the outer mold lines (OML) iges file. The internal duct was fabricated such that a linear transition occurs from the inlet to exhaust. Whereas, the iges definition has a constant area internal duct that quickly transitions from the inlet to exhaust cross sectional shape. The nacelle internal duct was fabricated, the way described, to save time and money. The variation in the cross sectional area is less than 1% from the iges definition. The nacelles were also installed with and without fairings. Fairings are defined as the build up of the nacelles on the upper wing surface so that the nacelles poke through the upper surface as defined in the OML iges file. Test 1679 data was of forces measured from model 2a and 2b. The wing for model 2b was a solid wing. The nacelles were built the same way as for model 2a, except for the nacelle base pressure installation. The nacelles were only tested with the fairings for model 2a and 2b during test 1679.

Cappuccio, Gelsomina↗

NASA NTRS2016

Statistically Based Approach to Broadband Liner Design and Assessment

A broadband liner design optimization includes utilizing in-duct attenuation predictions with a statistical fan source model to obtain optimum impedance spectra over a number of flow conditions for one or more liner locations in a bypass duct. The predicted optimum impedance information is then used with acoustic liner modeling tools to design liners having impedance spectra that most closely match the predicted optimum values. Design selection is based on an acceptance criterion that provides the ability to apply increasing weighting to specific frequencies and/or operating conditions. One or more broadband design approaches are utilized to produce a broadband liner that targets a full range of frequencies and operating conditions.

Nark, Douglas M.↗

NASA NTRS1991

The NASA Lewis Research Center Internal Fluid Mechanics Facility

An experimental facility specifically designed to investigate internal fluid duct flows is described. It is built in a modular fashion so that a variety of internal flow test hardware can be installed in the facility with minimal facility reconfiguration. The facility and test hardware interfaces are discussed along with design constraints of future test hardware. The plenum flow conditioning approach is also detailed. Available instrumentation and data acquisition capabilities are discussed. The incoming flow quality was documented over the current facility operating range. The incoming flow produces well behaved turbulent boundary layers with a uniform core. For the calibration duct used, the boundary layers approached 10 percent of the duct radius. Freestream turbulence levels at the various operating conditions varied from 0.64 to 0.69 percent of the average freestream velocity.

Porro, A. R.↗

NASA NTRS1993

CFD Modeling of Turbulent Duct Flows for Coolant Channel Analysis

The design of modern liquid rocket engines requires the analysis of chamber coolant channels to maximize the heat transfer while minimizing the coolant flow. Coolant channels often do not remain at a constant cross section or at uniform curvature. New designs require higher aspect ratio coolant channels than previously used. To broaden the analysis capability and to complement standard analysis tools an investigation on the accuracy of CFD predictions for coolant channel flow has been initiated. Validation of CFD capabilities for coolant channel analysis will enhance the capabilities for optimizing design parameters without resorting to extensive experimental testing. The eventual goal is to use CFD to determine the flow fields of unique coolant channel designs and therefore determine critical heat transfer coefficients. In this presentation the accuracy of a particular CFD code is evaluated for turbulent flows. The first part of the presentation is a comparison of numerical results to existing cold flow data for square curved ducts (NASA CR-3367, 'Measurements of Laminar and Turbulent Flow in a Curved Duct with Thin Inlet Boundary Layers'). The results of this comparison show good agreement with the relatively coarse experimental data. The second part of the presentation compares two cases of higher aspect ratio channels (AR=2.5,10) to show changes in axial and secondary flow strength. These cases match experimental work presently in progress and will be used for future validation. The comparison shows increased secondary flow strength of the higher aspect ratio case due to the change in radius of curvature. The presentation includes a test case with a heated wall to demonstrate the program's capability. The presentation concludes with an outline of the procedure used to validate the CFD code for future design analysis.

Ungewitter, Ronald J.↗

NASA NTRS1995

High frequency flow-structural interaction in dense subsonic fluids

Prediction of the detailed dynamic behavior in rocket propellant feed systems and engines and other such high-energy fluid systems requires precise analysis to assure structural performance. Designs sometimes require placement of bluff bodies in a flow passage. Additionally, there are flexibilities in ducts, liners, and piping systems. A design handbook and interactive data base have been developed for assessing flow/structural interactions to be used as a tool in design and development, to evaluate applicable geometries before problems develop, or to eliminate or minimize problems with existing hardware. This is a compilation of analytical/empirical data and techniques to evaluate detailed dynamic characteristics of both the fluid and structures. These techniques have direct applicability to rocket engine internal flow passages, hot gas drive systems, and vehicle propellant feed systems. Organization of the handbook is by basic geometries for estimating Strouhal numbers, added mass effects, mode shapes for various end constraints, critical onset flow conditions, and possible structural response amplitudes. Emphasis is on dense fluids and high structural loading potential for fatigue at low subsonic flow speeds where high-frequency excitations are possible. Avoidance and corrective measure illustrations are presented together with analytical curve fits for predictions compiled from a comprehensive data base.

Liu, Baw-Lin↗

NASA NTRS1975

Optimization of duct acoustic liners of finite length

Lining design procedures generally ignore the effects of discontinuous impedance changes on reflection and modal redistribution of incident energy. The present study involves identification of the mechanisms of optimal suppression by segmented liners, including impedance discontinuity effects. The mathematical model involves rectangular ducts in the absence of mean flow. Segmented acoustic fields are represented by duct eigenfunction expansions, and mode matching is used to ensure continuity of the total field. Extensive parameter studies were used to identify attenuation mechanisms and to define preliminary liner configurations. These are used as starting values in conjugate-gradient optimization studies. Optimal segmented liners are presented and it is shown that, if the sound source is sufficiently well-defined, conventional infinite duct optimum attenuation rates can be improved upon by taking advantage of wall impedance discontinuities. Finally, follow-on work which will involve experimental verification and extension to the case with mean flow in the duct is described.

Beckemeyer, R. J.↗

NASA NTRS1994

Improving Diffusing S-duct Performance by Secondary Flow Control

The objective of this research was to study ways to reduce inlet flow distortion (i.e., total pressure nonuniformity) and improve total pressure recovery in a diffusing S-duct. This was accomplished by controlling the development of secondary flows within the duct through the use of tapered-fin type vortex generators. Reported are results for the bare duct and seven different configurations of vortex generators. Data presented for each configuration include surface static pressure, surface flow visualization, and exit plane total pressure and transverse velocity. The performance of each configuration was assessed by calculating total pressure recovery and inlet distortion descriptors from the data and comparing them to the values for the bare duct. The best configuration tested reduced distortion (as measured by the DC(45) and DC(90) descriptors) by more than 50 percent while improving total pressure recovery by 0.5 percent. These results should provide valuable guidance in designing vortex generator installations in ducts and for assessing the accuracy of computational fluid dynamics (CFD) methods to calculate duct flows with installed vortex generators.

Reichert, Bruce A.↗

NASA NTRS1997

The NASA Lewis Research Center Water Tunnel Facility

A water tunnel facility specifically designed to investigate internal fluid duct flows has been built at the NASA Research Center. It is built in a modular fashion so that a variety of internal flow test hardware can be installed in the facility with minimal facility reconfiguration. The facility and test hardware interfaces are discussed along with design constraints for future test hardware. The inlet chamber flow conditioning approach is also detailed. Instrumentation and data acquisition capabilities are discussed. The incoming flow quality has been documented for about one quarter of the current facility operating range. At that range, there is some scatter in the data in the turbulent boundary layer which approaches 10 percent of the duct radius leading to a uniform core.

Wasserbauer, Charles A.↗

NASA NTRS1980

Thermodynamic performance testing of the orbiter flash evaporator system

System level testing of the space shuttle orbiter's development flash evaporator system (FES) was performed in a thermal vacuum chamber capable of simulating ambient ascent, orbital, and entry temperature and pressure profiles. The test article included the evaporator assembly, high load and topping exhaust duct and nozzle assemblies, and feedwater supply assembly. Steady state and transient heat load, water pressure/temperature and ambient pressure/temperature profiles were imposed by especially designed supporting test hardware. Testing in 1978 verified evaporator and duct heater thermal design, determined FES performance boundaries, and assessed topping evaporator plume characteristics. Testing in 1979 combined the FES with the other systems in the orbiter active thermal control subsystem (ATCS). The FES met or exceeded all nominal and contingency performance requirements during operation with the integrated ATCS. During both tests stability problems were encountered during steady state operations which resulted in subsequent design changes to the water spray nozzle and valve plate assemblies.

Jaax, J. R.↗

NASA NTRS1982

Frequency and Time Domain Modeling of Acoustic Liner Boundary Conditions

As part of a research program directed at the acoustics of advanced subsonic propulsion systems undertaken at NASA Langley, Duke University was funded to develop a boundary condition model for bulk-reacting nacelle liners. The overall objective of the Langley program was to understand and predict noise from advanced subsonic transport engines and to develop related noise control technology. The overall technical areas included: fan and propeller source noise, acoustics of ducts and duct liners, interior noise, subjective acoustics, and systems noise prediction. The Duke effort was directed toward duct liner acoustics for the development of analytical methods to characterize liner behavior in both frequency domain and time domain. A review of duct acoustics and liner technology can be found in Reference [1]. At that time, NASA Langley was investigating the propulsion concept of an advanced ducted fan, with a large diameter housed inside a relatively short duct. Fan diameters in excess of ten feet were proposed. The lengths of both the inlet and exhaust portions of the duct were to be short, probably less than half the fan diameter. The nacelle itself would be relatively thin-walled for reasons of aerodynamic efficiency. The blade-passage frequency was expected to be less than I kHz, and very likely in the 200 to 300 Hz range. Because of the design constraints of a short duct, a thin nacelle, and long acoustic wavelengths, the application of effective liner technology would be especially challenging. One of the needs of the NASA Langley program was the capability to accurately and efficiently predict the behavior of the acoustic liner. The traditional point impedance method was not an adequate model for proposed liner designs. The method was too restrictive to represent bulk reacting liners and to allow for the characterization of many possible innovative liner concepts. In the research effort at Duke, an alternative method, initially developed to handle bulk reacting layers as described in Reference [2], was extended to apply to a broad range of liner types. This method included the effect of local gradients along the liner surface, and was particularly appropriate for situations with flow over the liner and grazing incidence acoustic fields. In order to utilize time domain computational methods to solve for the propfan acoustic field, corresponding liner boundary conditions were developed for time domain solutions rather than frequency domain solutions.

Bliss, Donald B.↗

NASA NTRS2012

Preliminary Aerodynamic Investigation of Fan Rotor Blade Morphing

Various new technologies currently under development may enable controlled blade shape variability, or so-called blade morphing, to be practically employed in aircraft engine fans and compressors in the foreseeable future. The current study is a relatively brief, preliminary computational fluid dynamics investigation aimed at partially demonstrating and quantifying the aerodynamic potential of fan rotor blade morphing. The investigation is intended to provide information useful for near-term planning, as well as aerodynamic solution data sets that can be subsequently analyzed using advanced acoustic diagnostic tools, for the purpose of making fan noise comparisons. Two existing fan system models serve as baselines for the investigation: the Advanced Ducted Propulsor fan with a design tip speed of 806 ft/sec and a pressure ratio of 1.294, and the Source Diagnostic Test fan with a design tip speed of 1215 ft/sec and a pressure ratio of 1.470. Both are 22-in. sub-scale, low-noise research fan/nacelle models that have undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. The study, restricted to fan rotor blade morphing only, involves a fairly simple blade morphing technique. Specifically, spanwise-linear variations in rotor blade-section setting angle are applied to alter the blade shape; that is, the blade is linearly retwisted from hub to tip. Aerodynamic performance comparisons are made between morphed-blade and corresponding baseline configurations on the basis of equal fan system thrust, where rotor rotational speed for the morphed-blade fan is varied to change the thrust level for that configuration. The results of the investigation confirm that rotor blade morphing could be a useful technology, with the potential to enable significant improvements in fan aerodynamic performance. Even though the study is very limited in scope and confined to simple geometric perturbations of two existing fan systems, the aerodynamic effectiveness of blade morphing is demonstrated by the configurations analyzed. In particular, for the Advanced Ducted Propulsor fan it is demonstrated that the performance levels of the original variable-pitch baseline design can be achieved using blade morphing instead of variable pitch, and for the Source Diagnostic Test fan the performance at important off-design operating points is substantially increased with blade morphing.

Tweedt, Daniel L.↗

NASA NTRSDate not supplied

GT2024-139262: Development of a Computational Fluid Dynamics Model for a High-Speed Centrifugal Compressor

Computational fluid dynamics (CFD) has become widely used in the design and analysis of turbomachinery components such as centrifugal compressors. However, CFD is only a limited representation of experimental cases and struggles to model complex flows or lack model details to increase computation speed. The High Efficiency Centrifugal Compressor (HECC) was designed by United Technologies Research Center for NASA to investigate the difficulties behind improving centrifugal compressor technology and to provide an open case to the turbomachinery research community. The CFD technology seen in previous numerical models of the compressor have improved since then. In this study, a high-fidelity computational model for the vaneless configuration of HECC operating at its design point was developed and compared to an equivalent experimental case. This model incorporated GPU versions of mesh generation code WAND and solver code LEO from AeroDynamic Solutions, Inc. Blade sections extracted from the solid model formed the basis of the impeller mesh, while the remainder of the stage was modeled as ducts. Standard day aerodynamic conditions, design rotational speed, design mass flow rate, and inlet flow angles served as additional inputs. A mesh sensitivity study was performed prior to using the model for analysis. The accuracy of the model’s predictions of impeller and stage performance were assessed by comparing the total pressure ratio, total temperature rise ratio, and adiabatic efficiency to the corresponding parameters observed in the experimental case. Results showed a general overprediction of compressor performance by the model. Predictions were more accurate for the impeller performance compared to those for the full compressor stage. Spanwise analyses showed that the model tended to follow flow behavior patterns exhibited in the experiment. Beyond this study, this model was used to examine HECC aerodynamics and performance in research following thereafter.

centrifugal compressor↗

NASA NTRS1986

An experimental evaluation of S-duct inlet-diffuser configurations for turboprop offset gearbox applications

A test program, utilizing a large scale model, was run in the NASA Lewis Research Center 10- by 10-ft wind tunnel to examine the influence on performance of design parameters of turboprop S-duct inlet/diffuser systems. The parametric test program investigated inlet lip thickness, inlet/diffuser cross-sectional geometry, throat design Mach number, and shaft fairing shape. The test program was run at angles of attack to 15 deg and tunnel Mach numbers to 0.35. Results of the program indicate that current design techniques can be used to design inlet/diffuser systems with acceptable total pressure recovery, but several of the design parameters, notably lip thickness (contraction ratio) and shaft fairing cross section, must be optimized to prevent excessive distortion at the compressor face.

Mcdill, Paul L.↗

NASA NTRS2019

Numerical Investigation and Optimization of a Flushwall Injector for Scramjet Applications at Hypervelocity Flow Conditions

An investigation utilizing Reynolds-averaged simulations (RAS) was performed in order to demonstrate the use of design and analysis of computer experiments (DACE) methods in Sandia’s DAKOTA software package for surrogate modeling and optimization. These methods were applied to a flow- path fueled with an interdigitated flushwall injector suitable for scramjet applications at hyper- velocity conditions and ascending along a constant dynamic pressure flight trajectory. The flight Mach number, duct height, spanwise width, and injection angle were the design variables selected to maximize two objective functions: the thrust potential and combustion efficiency. Because the RAS of this case are computationally expensive, surrogate models are used for optimization. To build a surrogate model a RAS database is created. The sequence of the design variables comprising the database were generated using a Latin hypercube sampling (LHS) method. A methodology was also developed to automatically build geometries and generate structured grids for each design point. The ensuing RAS analysis generated the simulation database from which the two objective functions were computed using a one-dimensionalization (1D) of the three-dimensional simulation data. The data were fitted using four surrogate models: an artificial neural network (ANN), a cubic polynomial, a quadratic polynomial, and a Kriging model. Variance-based decomposition showed that both objective functions were primarily driven by changes in the duct height. Multiobjective design optimization was performed for all four surrogate models via a genetic algorithm method. Optimal solutions were obtained at the upper and lower bounds of the flight Mach number range. The Kriging model predicted an optimal solution set that exhibited high values for both objective functions. Additionally, three challenge points were selected to assess the designs on the Pareto fronts. Further sampling among the designs of the Pareto fronts may be required to lower the surrogate model errors and perform more accurate surrogate-model-based optimization.

Shenoy, Rajiv R.↗

NASA NTRS1999

ECLSS Design for the International Space Station Nodes 2 and 3

The International Space Station (ISS) modules Nodes 2 and 3 are currently under development by Alenia Spazio and the Marshall Space Flight Center (MSFC). The design of the Environmental Control and Life Support Systems (ECLSS) for these two modules have some similarities but many differences. The Node 2 ECLSS provides inter- and intramodule ventilation, temperature and humidity control, fire detection and suppression, and distribution of atmosphere samples, low pressure and recharge oxygen and nitrogen, fuel cell water and wastewater. Design Review 1 was held in March 1998. Fabrication of the ducting, tubing, and support structure is ongoing with Design Review 2 planned for December 1999. In addition to providing the same functions as Node 2, the ECLSS for Node 3 includes carbon dioxide removal, trace contaminant control, atmosphere monitoring, atmosphere pressure control, oxygen generation, urine and potable water processing, waste management, and potable water distribution to support the habitability functions relocated from the U.S. Hab module. Scarring for future incorporation of carbon dioxide reduction is also included. Design Review 1 is planned for July 1999. An overview of each Node's system design as well as issues and challenges are discussed.

Carrasquillo, Robyn L.↗