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At least 1,765 records · Page 98
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NASA NTRS2014

Wedge Shock and Nozzle Exhaust Plume Interaction in a Supersonic Jet Flow

Fundamental research for sonic boom reduction is needed to quantify the interaction of shock waves generated from the aircraft wing or tail surfaces with the nozzle exhaust plume. Aft body shock waves that interact with the exhaust plume contribute to the near-field pressure signature of a vehicle. The plume and shock interaction was studied using computational fluid dynamics and compared with experimental data from a coaxial convergent-divergent nozzle flow in an open jet facility. A simple diamond-shaped wedge was used to generate the shock in the outer flow to study its impact on the inner jet flow. Results show that the compression from the wedge deflects the nozzle plume and shocks form on the opposite plume boundary. The sonic boom pressure signature of the nozzle exhaust plume was modified by the presence of the wedge. Both the experimental results and computational predictions show changes in plume deflection.

Castner, Raymond↗
NASA NTRS2013

Rocket Sled Propelled Testing of a Supersonic Inflatable Aerodynamic Decelerator

Decelerators (IADs) have traditionally been tested in wind tunnels. As the limitations of these test facilities are reached, other avenues must be pursued. The IAD being tested is a Supersonic IAD (SIAD), which attaches just aft of the heatshield around the perimeter of an entry body. This 'attached torus' SIAD is meant to improve the accuracy of landing for robotic class missions to Mars and allow for potentially increased payloads. The SIAD Design Verification (SDV) test aims to qualify the SIAD by applying a targeted aerodynamic load to the vehicle. While many test architectures were researched, a rocket sled track was ultimately chosen to be the most cost effective way to achieve the desired dynamic pressures. The Supersonic Naval Ordnance Research Track (SNORT) at the Naval Air Warfare Center Weapons Division (NAWCWD) China Lake is a four mile test track, traditionally used for warhead and ejection seat testing. Prior to SDV, inflatable drag bodies have been tested on this particular track. Teams at Jet Propulsion Laboratory (JPL) and NAWCWD collaborate together to design and fabricate one of the largest sleds ever built. The SDV sled is comprised of three individual sleds: a Pusher Sled which holds the solid booster rockets, an Item Sled which supports the test vehicle, and a Camera Sled that is pushed in front for in-situ footage and measurements. The JPL-designed Test Vehicle has a full-scale heatshield shape and contains all instrumentation and inflation systems necessary to inflate and test a SIAD. The first campaign that is run at SNORT tested all hardware and instrumentation before the SIAD was ready to be tested. For each of the three tests in this campaign, the number of rockets and top speed was increased and the data analyzed to ensure the hardware is safe at the necessary accelerations and aerodynamic loads.

Supersonic Inflatable Aerodynamic Decelerators (SI↗
NASA NTRS2012

An Assessment of Open Rotor Noise Prediction Tools

Assess the current capability for predicting the aerodynamic and acoustic performance of open rotors. The testbed is a GE blade set called F31/A31 for which significant amount of aerodynamic and acoustic data was acquired in model scale tests. F31/A31 is a vintage 1990s design with a 12-bladed front rotor and a 10-bladed aft rotor. This blade set was tested in both low-speed regime (representative of approach and takeoff conditions) and high-speed regime (representative of climb and cruise conditions). Uninstalled as well as installed configurations were tested. The focus of this interim presentation is on a subset of the low-speed tests for which the tip speed was varied, but the blade setting angles and tunnel Mach number were held fixed.

Envia, Ed↗
NASA NTRS2015

VIIRS-J1 Polarization Narrative

The VIS/NIR bands polarization sensitivity of Joint Polar Satellite Sensor 1 (JPSS1) Visible/Infrared Imaging Radiometer Suite (VIIRS) instrument was measured using a broadband source. While polarization sensitivity for bands M5-M7, I1, and I2 was less than 2.5%, the maximum polarization sensitivity for bands M1, M2, M3, and M4 was measured to be 6.4%, 4.4%, 3.1%, and 4.3%, respectively with a polarization characterization uncertainty of less than 0.3%. A detailed polarization model indicated that the large polarization sensitivity observed in the M1 to M4 bands was mainly due to the large polarization sensitivity introduced at the leading and trailing edges of the newly manufactured VISNIR bandpass focal plane filters installed in front of the VISNIR detectors. This was confirmed by polarization measurements of bands M1 and M4 bands using monochromatic light. Discussed are the activities leading up to and including the instruments two polarization tests, some discussion of the polarization model and the model results, the role of the focal plane filters, the polarization testing of the Aft-Optics-Assembly, the testing of the polarizers at Goddard and NIST and the use of NIST's T-SIRCUS for polarization testing and associated analyses and results.

polarization↗
NASA NTRS2015

Liquid Hydrogen Propellant Tank Sub-Surface Pressurization with Gaseous Helium

A series of tests were conducted to evaluate the performance of a propellant tank pressurization system with the pressurant diffuser intentionally submerged beneath the surface of the liquid. Propellant tanks and pressurization systems are typically designed with the diffuser positioned to apply pressurant gas directly into the tank ullage space when the liquid propellant is settled. Space vehicles, and potentially propellant depots, may need to conduct tank pressurization operations in micro‐gravity environments where the exact location of the liquid relative to the diffuser is not well understood. If the diffuser is positioned to supply pressurant gas directly to the tank ullage space when the propellant is settled, then it may become partially or completely submerged when the liquid becomes unsettled in a microgravity environment. In such case, the pressurization system performance will be adversely affected requiring additional pressurant mass and longer pressurization times. This series of tests compares and evaluates pressurization system performance using the conventional method of supplying pressurant gas directly to the propellant tank ullage, and then supplying pressurant gas beneath the liquid surface. The pressurization tests were conducted on the Engineering Development Unit (EDU) located at Test Stand 300 at NASA Marshall Space Flight Center (MSFC). EDU is a ground based Cryogenic Fluid Management (CFM) test article supported by Glenn Research Center (GRC) and MSFC. A 150 ft3 propellant tank was filled with liquid hydrogen (LH2). The pressurization system used regulated ambient helium (GHe) as a pressurant, a variable position valve to maintain flow rate, and two identical independent pressurant diffusers. The ullage diffuser was located in the forward end of the tank and was completely exposed to the tank ullage. The submerged diffuser was located in the aft end of the tank and was completely submerged when the tank liquid level was 10% or greater. The ullage diffuser tests were conducted as a baseline to evaluate the performance of the pressurization system, and the submerged diffuser tests showed how the performance of the pressurization system was compromised when the diffuser was submerged in LH2. The test results are evaluated and compared, and included in this report for various propellant tank fill levels.

Stephens, J. R.↗
NASA NTRS2015

Novel Control Effectors for Truss Braced Wing

At cruise flight conditions very high aspect ratio/low sweep truss braced wings (TBW) may be subject to design requirements that distinguish them from more highly swept cantilevered wings. High aspect ratio, short chord length and relative thinness of the airfoil sections all contribute to relatively low wing torsional stiffness. This may lead to aeroelastic issues such as aileron reversal and low flutter margins. In order to counteract these issues, high aspect ratio/low sweep wings may need to carry additional high speed control effectors to operate when outboard ailerons are in reversal and/or must carry additional structural weight to enhance torsional stiffness. The novel control effector evaluated in this study is a variable sweep raked wing tip with an aileron control surface. Forward sweep of the tip allows the aileron to align closely with the torsional axis of the wing and operate in a conventional fashion. Aft sweep of the tip creates a large moment arm from the aileron to the wing torsional axis greatly enhancing aileron reversal. The novelty comes from using this enhanced and controllable aileron reversal effect to provide roll control authority by acting as a servo tab and providing roll control through intentional twist of the wing. In this case the reduced torsional stiffness of the wing becomes an advantage to be exploited. The study results show that the novel control effector concept does provide roll control as described, but only for a restricted class of TBW aircraft configurations. For the configuration studied (long range, dual aisle, Mach 0.85 cruise) the novel control effector provides significant benefits including up to 12% reduction in fuel burn.

White, Edward V.↗
NASA NTRS2015

Active Dihedral Control System for a Torisionally Flexible Wing

A span-loaded, highly flexible flying wing, having horizontal control surfaces mounted aft of the wing on extended beams to form local pitch-control devices. Each of five spanwise wing segments of the wing has one or more motors and photovoltaic arrays, and produces its own lift independent of the other wing segments, to minimize inter-segment loads. Wing dihedral is controlled by separately controlling the local pitch-control devices consisting of a control surface on a boom, such that inboard and outboard wing segment pitch changes relative to each other, and thus relative inboard and outboard lift is varied.

Kendall, Greg T.↗
NASA NTRS1959

Welded Titanium Case for Space-Probe Rocket Motor

Early in 1958, the Jet Propulsion Laboratory of the California Institute of Technology was requested to participate in a lunar-probe mission code-named Juno II which would place a 15-lb Instrumented payload (Pioneer IV) in the vicinity of the moon. The vehicle was to use the same high-speed upper-stage assembly as flown on the successful Jupiter-C configuration; however, the first-stage booster was to be a Jupiter rather than a Redstone. An analysis of the intended flight and payload configuration Indicated that the feasibility of accomplishing the mission was questionable and that additional performance would have to be obtained if the mission was to be feasible. Since the most efficient way of Increasing the performance of a staged vehicle is to increase the performance of the last stage, a study of possible ways of doing this was made.. Because of the time schedule placed on this effort It was decided to reduce the weight of the fourth-stage rocket-motor case by substituting the annealed 6Al--4V titanium alloy for the Type 410 stainless steel. Although this introduced an unfamiliar material, It reduced the changes in design and fabrication techniques. This particular titanium alloy was chosen on the basis of previous tests which proved the suitability of the alloy as a pressure-vessel material when used at an annealed yield strength of about 120, 000 psi. The titanium-case fourth stage of Juno U is shown with the payload and on the missile in Fig. 1; the stainless-steel motor cases used in the Jupiter-C vehicle are shown in Fig. 2. The fourth-stage motor case has a diameter of 6 in., a length of approximately 38 in. center dot and a nominal cylindrical wall thickness of 0.025 in. As shown in Fig. 1, the case serves as the structural support of the payload and is aligned to the upper stage assembly through an alignment ring. The nozzle is threaded into the end of the motor case, and is of the ceramic-coated steel design. Figure 3 shows a comparison of the components used to make the stainless steel and the 6A1--4V titanium alloy cases. The forward dome and aft fitting for the stainless steel assembly were fabricated from a combination of forged, spun and machined parts.. In order to facilitate the fabrication of the titanium alloy motor ) these components were machined from a large-diameter billet.

Brothers, A. J.↗
NASA NTRS2015

Observing System Simulations for Small Satellite Formations Estimating Bidirectional Reflectance

The bidirectional reflectance distribution function (BRDF) gives the reflectance of a target as a function of illumination geometry and viewing geometry, hence carries information about the anisotropy of the surface. BRDF is needed in remote sensing for the correction of view and illumination angle effects (for example in image standardization and mosaicing), for deriving albedo, for land cover classification, for cloud detection, for atmospheric correction, and other applications. However, current spaceborne instruments provide sparse angular sampling of BRDF and airborne instruments are limited in the spatial and temporal coverage. To fill the gaps in angular coverage within spatial, spectral and temporal requirements, we propose a new measurement technique: Use of small satellites in formation flight, each satellite with a VNIR (visible and near infrared) imaging spectrometer, to make multi-spectral, near-simultaneous measurements of every ground spot in the swath at multiple angles. This paper describes an observing system simulation experiment (OSSE) to evaluate the proposed concept and select the optimal formation architecture that minimizes BRDF uncertainties. The variables of the OSSE are identified; number of satellites, measurement spread in the view zenith and relative azimuth with respect to solar plane, solar zenith angle, BRDF models and wavelength of reflection. Analyzing the sensitivity of BRDF estimation errors to the variables allow simplification of the OSSE, to enable its use to rapidly evaluate formation architectures. A 6-satellite formation is shown to produce lower BRDF estimation errors, purely in terms of angular sampling as evaluated by the OSSE, than a single spacecraft with 9 forward-aft sensors. We demonstrate the ability to use OSSEs to design small satellite formations as complements to flagship mission data. The formations can fill angular sampling gaps and enable better BRDF products than currently possible.

small satellite↗
NASA NTRS1944

The Effect of Several Armament Installations on the Drag of 1/8-Scale Model of the B-32 Airplane

The effect of several armament installations on the drag of a 1/8 scale model of the B-32 airplane was determined. Turrets in the following positions were tested: nose, tail, upper forward, upper aft, and lower. The nose and tail turrets were each equipped with two.50-caliber guns. Upper turrets were of three types: two.50-caliber guns, four.50-caliber guns, and 20-millimeter cannon. Lower turrets were of two types: two.50-caliber guns and four.50-caliber guns. The effect of streamlining the upper two- and four-gun turrets and of extending the lower two-gun turret was determined. The tests were conducted in the Langley 19-foot. pressure tunnel at a Reynolds number of approximately 2,960,000 and a Mach number of 0.13. Large increases in drag coefficient were caused by the complete armament installations. At a lift coefficient of 0. 4 the installations with nonstreamlined upper turrets and the lower turret retracted increased the drag coefficient by 0.0022 and 0.0027 for the two-gun and four-gun turret installations, respectively. Streamlining the upper turrets reduced the drag of these installations by approximately 40 percent with the upper turrets streamlined, the drag increase was about the same for either the two- or four-gun turret installation. The streamlined two-cannon upper turrets increased the drag about the same amount as the two-gun upper turrets that were not streamlined. Extension of the lower turret. increased the drag slightly more than the whole streamlined gun-turret installation.

Fairbanks, R. W.↗
NASA NTRS2015

Evaluation of the Accuracy of Astroskin as a Behavioral Health Self-Monitoring System for Spaceflight

In space, there is a need to monitor astronauts' vital signs and assess their readiness to perform specific tasks during a mission. Currently, NASA does not have the capability to noninvasively monitor crew for extended periods of time. The Canadian Space Agency is working with the Psychophysiology Lab at NASA ARC to determine if the Astroskin could be used as a solution to this problem. Astroskin, a commercially available garment with built-in biosensors, can be comfortably worn under clothing or a spacesuit and relay information to the crewman's own mobile device. Data can also be sent wirelessly to the on-board Exploration Medical System. To determine if Astroskin meets requirements for health monitoring, it must first be validated in spaceflight analog environments. In the current study Astroskin data will be compared to traditional biomedical instrument measures of electrocardiography (ECG), respiration rate, and systolic blood pressure. The data will be recorded during Autogenic Feedback Training Exercise (AFTE), which is a type of physiological self-regulation training designed for astronauts. The data will also be recorded during simulations of the Orion spacecraft re-entry. The results to date suggest that Astroskin is a suitable ambulatory monitoring system that allows astronauts to self-diagnose and self-regulate adverse autonomic nervous system responses to sustained exposure to microgravity of spaceflight.

biomedical↗
NASA NTRS2015

Jet-Surface Interaction: High Aspect Ratio Nozzle Test, Nozzle Design and Preliminary Data

The Jet-Surface Interaction High Aspect Ratio (JSI-HAR) nozzle test is part of an ongoing effort to measure and predict the noise created when an aircraft engine exhausts close to an airframe surface. The JSI-HAR test is focused on parameters derived from the Turbo-electric Distributed Propulsion (TeDP) concept aircraft which include a high-aspect ratio mailslot exhaust nozzle, internal septa, and an aft deck. The size and mass flow rate limits of the test rig also limited the test nozzle to a 16:1 aspect ratio, half the approximately 32:1 on the TeDP concept. Also, unlike the aircraft, the test nozzle must transition from a single round duct on the High Flow Jet Exit Rig, located in the AeroAcoustic Propulsion Laboratory at the NASA Glenn Research Center, to the rectangular shape at the nozzle exit. A parametric nozzle design method was developed to design three low noise round-to-rectangular transitions, with 8:1, 12:1, and 16: aspect ratios, that minimizes flow separations and shocks while providing a flat flow profile at the nozzle exit. These designs validated using the WIND-US CFD code. A preliminary analysis of the test data shows that the actual flow profile is close to that predicted and that the noise results appear consistent with data from previous, smaller scale, tests. The JSI-HAR test is ongoing through October 2015. The results shown in the presentation are intended to provide an overview of the test and a first look at the preliminary results.

Aeroacoustics↗
NASA NTRS2012

Flowfield Analysis of a Small Entry Probe (SPRITE) Tested in an Arc Jet

A novel concept of small size (diameter less than 15 inches) entry probes named SPRITE (Small Probe Re-entry Investigation for TPS Engineering) has been developed at NASA Ames Research Center (ARC). These flight probes have on-board data acquisition systems that have also been developed in parallel at NASA ARC by Greg Swanson1. Flight probes of this size facilitate testing over a wide range of conditions in arc jets available at NASA ARC, thereby fulfilling a 'test what you fly' paradigm. As indicated by the acronym, these probes, with suitably tailored trajectories, are primarily meant to be robotic flight test beds for TPS materials, although the design is flexible enough to accommodate additional objectives of flight-testing other vehicle subsystems. A first step towards establishing the feasibility of the SPRITE concept is to arc-jet test fully instrumented models at flight scale. In a follow-on to the Large-Scale Article Tests (LSAT2) performed in the 60 MW Interaction Heating Facility (IHF) in late 2008/early 2009, a full-scale model of Deep Space-2 (DS23) made of red oak was tested in the 20 MW Aerodynamic Heating Facility (AHF). There were no issues with mass capture by the diffuser for blunt bodies of roughly 15 inches diameter tested in the 18-inch nozzle of the AHF. Building on this initial success, two identical test articles - SPRITE-T1-1 and SPRITE-T1-2 (T1 indicating the choice of back shell geometry) - were fabricated, and one of them, SPRITE-T1-1, was tested in the AHF recently. Both these test articles, 14 inches in diameter, have a 45deg sphere-cone (like DS2) made of PICA bonded on to a 1/8th inch thick aluminum shell using RTV. The aft portion of the test article is a conical frustum (15deg cone angle) with LI-2200 bonded on to the aluminum shell. Each model is fully instrumented with: (a) thermocouples imbedded in plugs in the heat shield, (b) thermocouples bonded to the aluminum substructure; the thermocouples are distributed over the entire shell, and (c) a few strain gages. Data from some of the thermocouples and gages are acquired by the on-board data acquisition system (DAS), while data from the others are routed to the facility-provided DAS, thereby enabling a cross check on the in situ measurement capability. as inputs to v2.6.1 of the in-house materials thermal response code, FIAT

atmospheric entry↗
NASA NTRS2016

Computational Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Takeoff and Landing Conditions

Due to the aft, upper surface engine location on the Hybrid Wing Body (HWB) planform, there is potential to shed vorticity and separated wakes into the engine when the vehicle is operated at off-design conditions and corners of the envelope required for engine and airplane certification. CFD simulations were performed of the full-scale reference propulsion system, operating at a range of inlet flow rates, flight speeds, altitudes, angles of attack, and angles of sideslip to identify the conditions which produce the largest distortion and lowest pressure recovery. Pretest CFD was performed by NASA and Boeing, using multiple CFD codes, with various turbulence models. These data were used to make decisions regarding model integration, characterize inlet flow distortion patterns, and help define the wind tunnel test matrix. CFD was also performed post-test; when compared with test data, it was possible to make comparisons between measured model-scale and predicted full-scale distortion levels. This paper summarizes these CFD analyses.

Wind Tunnel Testing↗
NASA NTRS2016

Pick-and-Eat Salad-Crop Productivity, Nutritional Value, and Acceptability to Supplement the ISS Food System

The capability to grow nutritious, palatable food for crew consumption during spaceflight has the potential to provide health-promoting, bioavailable nutrients, enhance the dietary experience, and reduce launch mass as we move toward longer-duration missions. However, studies of edible produce during spaceflight have been limited, leaving a significant knowledge gap in the methods required to grow safe, acceptable, nutritious crops for consumption in space. Researchers from Kennedy Space Center, Johnson Space Center, Purdue University and ORBITEC have teamed up to explore the potential for plant growth and food production on the International Space Station (ISS) and future exploration missions. KSC, Purdue, and ORBITEC bring a history of plant and plant-microbial interaction research for ISS and for future bioregenerative life support systems. JSC brings expertise in Advanced Food Technology (AFT), Behavioral Health and Performance (BHP), and statistics. The Veggie vegetable-production system on the ISS offers an opportunity to develop a pick-and-eat fresh vegetable component to the ISS food system as a first step to bioregenerative supplemental food production. We propose growing salad plants in the Veggie unit during spaceflight, focusing on the impact of light quality and fertilizer formulation on crop morphology, edible biomass yield, microbial food safety, organoleptic acceptability, nutritional value, and behavioral health benefits of the fresh produce. The first phase of the project will involve flight tests using leafy greens, with a small Chinese cabbage variety, Tokyo bekana, previously down selected through a series of research tests as a suitable candidate. The second phase will focus on dwarf tomato. Down selection of candidate varieties have been performed, and the dwarf cultivar Red Robin has been selected as the test crop. Four light treatments and three fertilizer treatments will be tested for each crop on the ground, to down select to two light treatments and one fertilizer treatment to test on ISS. Our work will help define light colors, levels, and horticultural best practices to achieve high yields of safe, nutritious leafy greens and tomatoes to supplement a space diet of prepackaged food. Our final deliverable will be the development of growth protocols for these crops in a spaceflight vegetable production system. With this work, and potentially with other pending joint projects, we will continue the synergistic research to help close gaps in the human research roadmap, and enable humans to venture to Mars and beyond. This research was co-funded by the Human Research Program and Space Biology (MTL1075) in the ILSRA 2015 NRA call.

Behavioral Health↗
NASA NTRS2016

Space Launch System Base Heating Test: Environments and Base Flow Physics

The NASA Space Launch System (SLS) vehicle is composed of four RS-25 liquid oxygen-hydrogen rocket engines in the core-stage and two 5-segment solid rocket boosters and as a result six hot supersonic plumes interact within the aft section of the vehicle during flight. Due to the complex nature of rocket plume-induced flows within the launch vehicle base during ascent and a new vehicle configuration, sub-scale wind tunnel testing is required to reduce SLS base convective environment uncertainty and design risk levels. This hot-fire test program was conducted at the CUBRC Large Energy National Shock (LENS) II short-duration test facility to simulate flight from altitudes of 50 kft to 210 kft. The test program is a challenging and innovative effort that has not been attempted in 40+ years for a NASA vehicle. This paper discusses the various trends of base convective heat flux and pressure as a function of altitude at various locations within the core-stage and booster base regions of the two-percent SLS wind tunnel model. In-depth understanding of the base flow physics is presented using the test data, infrared high-speed imaging and theory. The normalized test design environments are compared to various NASA semi-empirical numerical models to determine exceedance and conservatism of the flight scaled test-derived base design environments. Brief discussion of thermal impact to the launch vehicle base components is also presented.

Mehta, Manish↗
NASA NTRS2014

Stardust Interstellar Preliminary Examination IV: Scanning Transmission X-Ray Microscopy Analyses of Impact Features in the Stardust Interstellar Dust Collector

We report the quantitative characterization by synchrotron soft X-ray spectroscopy of 31 potential impact features in the aerogel capture medium of the Stardust Interstellar Dust Collector. Samples were analyzed in aerogel by acquiring high spatial resolution maps and high energy-resolution spectra of major rock-forming elements Mg, Al, Si, Fe, and others. We developed diagnostic screening tests to reject spacecraft secondary ejecta and terrestrial contaminants from further consideration as interstellar dust candidates. The results support an extraterrestrial origin for three interstellar candidates: I1043,1,30 (Orion) is a 3 pg particle with Mg-spinel, forsterite, and an iron-bearing phase. I1047,1,34 (Hylabrook) is a 4 pg particle comprising an olivine core surrounded by low-density, amorphous Mg-silicate and amorphous Fe, Cr, and Mn phases. I1003,1,40 (Sorok) has the track morphology of a high-speed impact, but contains no detectable residue that is convincingly distinguishable from the background aerogel. Twenty-two samples with an anthropogenic origin were rejected, including four secondary ejecta from impacts on the Stardust spacecraft aft solar panels, nine ejecta from secondary impacts on the Stardust Sample Return Capsule, and nine contaminants lacking evidence of an impact. Other samples in the collection included I1029,1,6, which contained surviving solar system impactor material. Four samples remained ambiguous: I1006,2,18, I1044,2,32, and I1092,2,38 were too dense for analysis, and we did not detect an intact projectile in I1044,3,33. We detected no radiation effects from the synchrotron soft X-ray analyses; however, we recorded the effects of synchrotron hard X-ray radiation on I1043,1,30 and I1047,1,34.

Dust↗
NASA NTRS2016

Computational Investigation of a Boundary-Layer Ingestion Propulsion System for the Common Research Model

This thesis will examine potential propulsive and aerodynamic benefits of integrating a boundary-layer ingestion (BLI) propulsion system with a typical commercial aircraft using the Common Research Model geometry and the NASA Tetrahedral Unstructured Software System (TetrUSS). The Numerical Propulsion System Simulation (NPSS) environment will be used to generate engine conditions for CFD analysis. Improvements to the BLI geometry will be made using the Constrained Direct Iterative Surface Curvature (CDISC) design method. Previous studies have shown reductions of up to 25% in terms of propulsive power required for cruise for other axisymmetric geometries using the BLI concept. An analysis of engine power requirements, drag, and lift coefficients using the baseline and BLI geometries coupled with the NPSS model are shown. Potential benefits of the BLI system relating to cruise propulsive power are quantified using a power balance method and a comparison to the baseline case is made. Iterations of the BLI geometric design are shown and any improvements between subsequent BLI designs presented. Simulations are conducted for a cruise flight condition of Mach 0.85 at an altitude of 38,500 feet and an angle of attack of 2deg for all geometries. A comparison between available wind tunnel data, previous computational results, and the original CRM model is presented for model verification purposes along with full results for BLI power savings. Results indicate a 14.3% reduction in engine power requirements at cruise for the BLI configuration over the baseline geometry. Minor shaping of the aft portion of the fuselage using CDISC has been shown to increase the benefit from boundary-layer ingestion further, resulting in a 15.6% reduction in power requirements for cruise as well as a drag reduction of eighteen counts over the baseline geometry.

Blumenthal, Brennan↗