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All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Active Flow Control (AFC) and Insect Accretion and Mitigation (IAM) System Design and Integration on the Boeing 757 ecoDemonstrator

This paper presents a systems overview of how the Boeing and NASA team designed, analyzed, fabricated, and integrated the Active Flow Control (AFC) technology and Insect Accretion Mitigation (IAM) systems on the Boeing 757 ecoDemonstrator. The NASA Environmentally Responsible Aviation (ERA) project partnered with Boeing to demonstrate these two technology systems on a specially outfitted Boeing 757 ecoDemonstrator during the spring of 2015. The AFC system demonstrated attenuation of flow separation on a highly deflected rudder and increased the side force generated. This AFC system may enable a smaller vertical tail to provide the control authority needed in the event of an engine failure during takeoff while still operating in a conventional manner over the rest of the flight envelope. The AFC system consisted of ducting to obtain air from the Auxiliary Power Unit (APU), a control valve to modulate the system mass flow, a heat exchanger to lower the APU air temperature, and additional ducting to deliver the air to the AFC actuators located on the vertical tail. The IAM system demonstrated how to mitigate insect residue adhesion on a wing's leading edge. Something as small as insect residue on a leading edge can cause turbulent wedges that interrupt laminar flow, resulting in an increase in drag and fuel use. The IAM system consisted of NASA developed Engineered Surfaces (ES) which were thin aluminum sheet substrate panels with coatings applied to the exterior. These ES were installed on slats 8 and 9 on the right wing of the 757 ecoDemonstrator. They were designed to support panel removal and installation in one crew shift. Each slat accommodated 4 panels. Both the AFC and IAM flight test were the culmination of several years of development and produced valuable data for the advancement of modern aircraft designs.

Alexander, Michael G.↗
NASA NTRS1973

On the Design and Test of a Liquid Injection Electric Thruster

A liquid injection electric thruster (LINJET) was designed and tested. The results of the tests were very encouraging with thruster performance levels well in excess of design goals. Supporting activities to the engine design and test included a five-million pulse life test on the main capacitor, a 46-million pulse test on the trigger electronics, design and fabrication of a zero resistance torque connector for use with the torsional pendulum thrust stand, design and fabrication of a logic box for control of engine firing, and a physical and chemical properties characterization of the perfluorocarbon propellant. While the results were encouraging, testing was limited, as many problems existed with the design. The most significant problem was involved with excessive propellant flow which contributed to false triggering and shorting. Low power active thermal control of the propellant storage cavity, coupled with a re-evaluation of the injection ring pore size and area exposed to the main capacitor discharge are areas that should be investigated should this design be carried forward.

Jones, T. A.↗
NASA NTRS1995

The NASA Langley 16-Foot Transonic Tunnel: Historical Overview, Facility Description, Calibration, Flow Characteristics, and Test Capabilities

The Langley 16-Foot Transonic Tunnel is a closed-circuit single-return atmospheric wind tunnel that has a slotted octagonal test section with continuous air exchange. The wind tunnel speed can be varied continuously over a Mach number range from 0.1 to 1.3. Test-section plenum suction is used for speeds above a Mach number of 1.05. Over a period of some 40 years, the wind tunnel has undergone many modifications. During the modifications completed in 1990, a new model support system that increased blockage, new fan blades, a catcher screen for the first set of turning vanes, and process controllers for tunnel speed, model attitude, and jet flow for powered models were installed. This report presents a complete description of the Langley 16-Foot Transonic Tunnel and auxiliary equipment, the calibration procedures, and the results of the 1977 and the 1990 wind tunnel calibration with test section air removal. Comparisons with previous calibrations showed that the modifications made to the wind tunnel had little or no effect on the aerodynamic characteristics of the tunnel. Information required for planning experimental investigations and the use of test hardware and model support systems is also provided.

Capone, Francis J.↗
NASA NTRS2014

Temperature Oscillation in a Loop Heat Pipe with Gravity Assist

ATLAS Laser Thermal Control System (LTCS) thermal vacuum testing where the condenser-radiator was placed in a vertical position, it was found that the loop heat pipe (LHP) reservoir required much more control heater power than the analytical model had predicted. The required control heater power was also higher than the liquid subcooling entering the reservoir using the measured temperatures and the calculated mass flow rate based on steady state LHP operation. This presentation describes the investigation of the LHP behaviors under a gravity assist mode with a very cold radiator sink temperature and a large thermal mass attached to the evaporator. It is concluded that gravity caused the cold liquid to drop from the condenser-radiator to the reservoir, resulting in a rapid decrease of the reservoir temperature. When the reservoir temperature was increasing, a reverse flow occurred in the liquid line, carrying warm liquid to the condenser-radiator. Both events consumed the reservoir control heater power. The fall and rise of the reservoir temperature also caused the net heat input to the evaporator to vary due to the release and storage of the sensible heat of the thermal mass. The combination of these effects led to a persistent reservoir temperature oscillation and a repeated influx of cold liquid from the condenser. This was the root cause of the extraordinary high control heater power requirement in the LTCS TV test. Without gravity assist, such a persistent temperature oscillation will not be present.

000↗
NASA NTRS2004

Overview of the GRC Stirling Convertor System Dynamic Model

A Stirling Convertor System Dynamic Model has been developed at the Glenn Research Center for controls, dynamics, and systems development of free-piston convertor power systems. It models the Stirling cycle thermodynamics, heat flow, gas, mechanical, and mounting dynamics, the linear alternator, and the controller. The model's scope extends from the thermal energy input to thermal, mechanical dynamics, and electrical energy out, allowing one to study complex system interactions among subsystems. The model is a non-linear time-domain model containing sub-cycle dynamics, allowing it to simulate transient and dynamic phenomena that other models cannot. The model details and capability are discussed.

Lewandowski, Edward J.↗
NASA NTRS2016

Digital Valve for High Pressure High Flow Applications

To address the challenges, which are involved with the development of flow control valves that can meet the requirements of deep oil wells such as high pressure, high flow rate, limited power and limited space, the authors have conceived a novel design configuration [1]. This design consists of a digitalized flow control valve with multipath and multistage pressure reduction structures. Specifically, the valve is configured as a set of parallel flow paths from the inlet to the outlet. A choke valve controls the total flow rate by digitally opening different paths or different combination of the paths. Each path is controlled by a poppet cap valve basically operated in on-off states. The number of flow states is 2N where N is the number of flow paths. To avoid erosion from sand in the oil and high speed flow, the seal area of the poppet cap valve is located at a distance from the flow inlet away from the high speed flow and the speed is controlled to stay below a predefined erosion safe limit. The path is a multistage structure composed of a set of serial nozzles-expansion chambers that equally distribute the total pressure drop to each stage. The pressure drop of each stage and, therefore, the flow speed at the nozzles and expansion chambers is controlled by the number of stages. The paths have relatively small cross section and could be relatively long for large number of stages and still fit in the strict annular space limit of a typical downhole region of an oil well. The paper will present the design configuration, analysis and preliminary test results.

downhole↗
NASA NTRS2016

Mechanisms of Active Aerodynamic Load Reduction on a Rotorcraft Fuselage With Rotor Effects

The reduction of the aerodynamic load that acts on a generic rotorcraft fuselage by the application of active flow control was investigated in a wind tunnel test conducted on an approximately 1/3-scale powered rotorcraft model simulating forward flight. The aerodynamic mechanisms that make these reductions, in both the drag and the download, possible were examined in detail through the use of the measured surface pressure distribution on the fuselage, velocity field measurements made in the wake directly behind the ramp of the fuselage and computational simulations. The fuselage tested was the ROBIN-mod7, which was equipped with a series of eight slots located on the ramp section through which flow control excitation was introduced. These slots were arranged in a U-shaped pattern located slightly downstream of the baseline separation line and parallel to it. The flow control excitation took the form of either synthetic jets, also known as zero-net-mass-flux blowing, and steady blowing. The same set of slots were used for both types of excitation. The differences between the two excitation types and between flow control excitation from different combinations of slots were examined. The flow control is shown to alter the size of the wake and its trajectory relative to the ramp and the tailboom and it is these changes to the wake that result in a reduction in the aerodynamic load.

Schaeffler, Norman W.↗
NASA NTRS1982

COSAL: A black-box compressible stability analysis code for transition prediction in three-dimensional boundary layers

A fast computer code COSAL for transition prediction in three dimensional boundary layers using compressible stability analysis is described. The compressible stability eigenvalue problem is solved using a finite difference method, and the code is a black box in the sense that no guess of the eigenvalue is required from the user. Several optimization procedures were incorporated into COSAL to calculate integrated growth rates (N factor) for transition correlation for swept and tapered laminar flow control wings using the well known e to the Nth power method. A user's guide to the program is provided.

Malik, M. R.↗
NASA NTRS2000

Airborne Measurements of Nitric Oxide, Nitrogen Dioxide, Ozone, and Total Reactive Nitrogen During the NASA Global Tropospheric Experiment

Fabrication of the University of Michigan Multichannel Chemiluminescence Instrument (UMMCI) was completed in early 1996 and the instrument participated in test flights on the NASA P3B at Wallops Island prior to integration and deployment for the PEM- Tropics A Mission. The UMMCI consists of 4 channels for simultaneous measurements of ozone and NO with the option for measurements of NO2 and NOy (total reactive nitrogen) when converters are placed upstream of the NO channels. Each NO channel consists of a zeroing volume and reaction vessel, while the ozone channel consists of an ozone catalyst (or scrubber) trap that is not in line with the reaction vessel. The detectors in all for channels are Hamamatsu photomultiplier tubes, which are followed by pulse amplifier discriminators on the NO channels and an electrometer on the ozone channel. Schematics of the Detector Module and NOx/03 Probe Insert and Diagrams of the Control and Data System, the Power and Ground System, the Gas Flow System, and the Calibration System Flow are attached. Intercomparisons were conducted with G. Gregory, NASA/Langley, during the test flights (following prior calibration of the ozone generator/calibrators at the Wallops Long-Path Absorption facility). Initial test results appeared to be reasonable, and instrument characterization studies proceeded for the ozone channel and the 3 NO channels until deployment for integration for the PEM-Tropics Mission. Ozone data was obtained for Flights #4, and 6-2 1, and finalized data was submitted to the PEM-Tropics Data Archive and to the Science Team during the April 1997 Data Workshop. Although it initially appeared that the instrument sensitivity varied, subsequent tests showed that this was the fault of a leak in the ozone calibrator. In fact; the instrument sensitivity has not been observed to vary in a large number of tests over the years since the PEM-Tropics mission. We have, therefore, a very high degree of confidence in the O3 data that we submitted. NO data was obtained for all flights except the mission out of Christmas Island and the subsequent return to Hawaii, during which time the NO channels were contaminated with back-flushing pure NO from the O3 channel. The NO channels were found to suffer from a varying artifact (7 - 22 pptv) which made the instrument's results unreliable for the marine boundary layer portions of the missions (where conditions were extremely clean with NO typically < 2 pptv and O3 only 8 ppbv). Nonetheless, the NO values were an upper limit and were quite useful in identifying pollution layers and outflow from the Latin American Continent. NO data was submitted to the archive for all flights during which data was obtained. Ozone data from the P3B was among the results presented in several peer-reviewed publications. Notably, the ozone data supported the first published finding of evidence for the occurrence of aerosol nucleation in the marine boundary layer. As well, the ozone data was utilized in a treatment of the wavenumber spectra of winds, temperature and trace gases and to demonstrate the occurrence of layering as observed by the P3B during PEM-Tropics A . O3 and NO data were used in an assessment of the impact of southern hemispheric biomass burning, and in the role of NOx and O3 in HOx photochemistry during the PEM Tropics A mission. These data, along with other species measured during PEM-Tropics A, have also been employed in the generation of climatologies for tropospheric ozone and its precursors by the PI's research group.

Carroll, Mary Anne↗
NASA NTRS2003

Magnetic Flux Compression Experiments Using Plasma Armatures

Magnetic flux compression reaction chambers offer considerable promise for controlling the plasma flow associated with various micronuclear/chemical pulse propulsion and power schemes, primarily because they avoid thermalization with wall structures and permit multicycle operation modes. The major physical effects of concern are the diffusion of magnetic flux into the rapidly expanding plasma cloud and the development of Rayleigh-Taylor instabilities at the plasma surface, both of which can severely degrade reactor efficiency and lead to plasma-wall impact. A physical parameter of critical importance to these underlying magnetohydrodynamic (MHD) processes is the magnetic Reynolds number (R(sub m), the value of which depends upon the product of plasma electrical conductivity and velocity. Efficient flux compression requires R(sub m) less than 1, and a thorough understanding of MHD phenomena at high magnetic Reynolds numbers is essential to the reliable design and operation of practical reactors. As a means of improving this understanding, a simplified laboratory experiment has been constructed in which the plasma jet ejected from an ablative pulse plasma gun is used to investigate plasma armature interaction with magnetic fields. As a prelude to intensive study, exploratory experiments were carried out to quantify the magnetic Reynolds number characteristics of the plasma jet source. Jet velocity was deduced from time-of-flight measurements using optical probes, and electrical conductivity was measured using an inductive probing technique. Using air at 27-inHg vacuum, measured velocities approached 4.5 km/s and measured conductivities were in the range of 30 to 40 kS/m.

Turner, M. W.↗
NASA NTRS2005

Heat Pipe Reactor Dynamic Response Tests: SAFE-100 Reactor Core Prototype

The SAFE-I00a test article at the NASA Marshall Space Flight Center was used to simulate a variety of potential reactor transients; the SAFEl00a is a resistively heated, stainless-steel heat-pipe (HP)-reactor core segment, coupled to a gas-flow heat exchanger (HX). For these transients the core power was controlled by a point kinetics model with reactivity feedback based on core average temperature; the neutron generation time and the temperature feedback coefficient are provided as model inputs. This type of non-nuclear test is expected to provide reasonable approximation of reactor transient behavior because reactivity feedback is very simple in a compact fast reactor (simple, negative, and relatively monotonic temperature feedback, caused mostly by thermal expansion) and calculations show there are no significant reactivity effects associated with fluid in the HP (the worth of the entire inventory of Na in the core is .<el, so fluid movement and temperature changes will cause very minor effects). In previous SAFE-100 tests, the point kinetics model was based on core thermal expansion via deflection measurements. It was found that core deflection was a strung function of how the SAFE-100 modules were fabricated and assembled (in terms of straightness, gaps, and other tolerances). To remove the added variable of how this particular core expands as compared to a different concept, it was decided to use a temperature based feedback model (based on several thermocouples placed throughout the core).

Bragg-Sitton, Shannon M.↗
NASA NTRS2004

Turbofan Engine Simulated in a Graphical Simulation Environment

Recently, there has been an increase in the development of intelligent engine technology with advanced active component control. The computer engine models used in these control studies are component-level models (CLM), models that link individual component models of state space and nonlinear algebraic equations, written in a computer language such as Fortran. The difficulty faced in performing control studies on Fortran-based models is that Fortran is not supported with control design and analysis tools, so there is no means for implementing real-time control. It is desirable to have a simulation environment that is straightforward, has modular graphical components, and allows easy access to health, control, and engine parameters through a graphical user interface. Such a tool should also provide the ability to convert a control design into real-time code, helping to make it an extremely powerful tool in control and diagnostic system development. Simulation time management is shown: Mach number versus time, power level angle versus time, altitude versus time, ambient temperature change versus time, afterburner fuel flow versus time, controller and actuator dynamics, collect initial conditions, CAD output, and component-level model: CLM sensor, CAD input, and model output. The Controls and Dynamics Technologies Branch at the NASA Glenn Research Center has developed and demonstrated a flexible, generic turbofan engine simulation platform that can meet these objectives, known as the Modular Aero-Propulsion System Simulation (MAPSS). MAPSS is a Simulink-based implementation of a Fortran-based, modern high pressure ratio, dual-spool, low-bypass, military-type variable-cycle engine with a digital controller. Simulink (The Mathworks, Natick, MA) is a computer-aided control design and simulation package allows the graphical representation of dynamic systems in a block diagram form. MAPSS is a nonlinear, non-real-time system composed of controller and actuator dynamics (CAD) and component-level model (CLM) modules. The controller in the CAD module emulates the functionality of a digital controller, which has a typical update rate of 50 Hz. The CLM module simulates the dynamics of the engine components and uses an update rate of 2500 Hz, which is needed to iterate to balance mass and energy among system components. The actuators in the CAD module use the same sampling rate as those in the CLM. Two graphs of normalized spool speed versus time in seconds and one graph of normalized average metal temperature versus time in seconds is shown. MAPSS was validated via open-loop and closed-loop comparisons with the Fortran simulation. The preceding plots show the normalized results of a closed-loop comparison looking at three states of the model: low-pressure spool speed, high-pressure spool speed, and the average metal temperature measured from the combustor to the high-pressure turbine. In steady state, the error between the simulations is less than 1 percent. During a transient, the difference between the simulations is due to a correction in MAPSS that prevents the gas flow in the bypass duct inlet from flowing forward instead of toward the aft end, which occurs in the Fortran simulation. A comparison between MAPSS and the Fortran model of the bypass duct inlet flow for power lever angles greater than 35 degrees is shown.

Parker, Khary I.↗
NASA NTRS1992

Using artificial intelligence to control fluid flow computations

Computational simulation is an essential tool for the prediction of fluid flow. Many powerful simulation programs exist today. However, using these programs to reliably analyze fluid flow and other physical situations requires considerable human effort and expertise to set up a simulation, determine whether the output makes sense, and repeatedly run the simulation with different inputs until a satisfactory result is achieved. Automating this process is not only of considerable practical importance but will also significantly advance basic artificial intelligence (AI) research in reasoning about the physical world.

Gelsey, Andrew↗
NASA NTRS1975

Use of ERTS-1 DCS in the management and control of water resources systems

The ERTS-1 experimental hydrologic Data Collection Platform System that has been established at the New England Division (NED), the reasons for getting involved with the experiment, some of the initial problems associated with the data collection hardware, and a preliminary conclusion based on operating experiences are reviewed. The New England Region includes the states of Maine, New Hampshire, Vermont, Massachusetts, Rhode Island and Connecticut. The entire area consists of approximately 97,000 sq. km. (60,000 square miles), half of which is in the state of Maine. The limits of the NED are all of Maine, New Hampshire and Vermont to the western limits of the Connecticut River basin, Massachusetts, Connecticut to the western edge of the Housatonic River basin and Rhode Island. All reservoirs have flood control as a primary purpose. Other uses include water supply, recreation and low flow augmentation. However, none of the reservoirs are presently operated for hydroelectric power, navigation, or irrigation purposes. Basically then, flood control regulation is NED's primary concern.

Finegan, J. W., Jr.↗
NASA NTRS2003

Experimental Evaluation of Hot Films on Ceramic Substrates for Skin-Friction Measurement

An investigation has been performed on the use of low-thermal conductivity, ceramic substrates for hot films intended to measure skin friction. Hot films were deposited on two types of ceramic substrates. Four hot films used composite-ceramic substrates with subsurface thermocouples (TCs), and two hot films were deposited on thin Macor(R) substrates. All six sensors were tested side by side in the wall of the NASA Glenn Research Center 8-ft by 6-ft Supersonic Wind Tunnel (SWT). Data were obtained from zero flow to Mach 1.98 in air. Control measurements were made with three Preston tubes and two boundary-layer rakes. The tests were repeated at two different hot film power levels. All hot films and subsurface TCs functioned throughout the three days of testing. At zero flow, the films on the high-thermal conductivity Macor(R) substrates required approximately twice the power as those on the composite-ceramic substrates. Skin-friction results were consistent with the control measurements. Estimates of the conduction heat losses were made using the embedded TCs but were hampered by variability in coating thicknesses and TC locations.

Noffz, Gregory K.↗
NASA NTRS2015

Low-Power Operation and Plasma Characterization of a Qualification Model SPT-140 Hall Thruster for NASA Science Missions

Hall thruster systems based on commercial product lines can potentially lead to lower cost electric propulsion (EP) systems for deep space science missions. A 4.5-kW SPT-140 Hall thruster presently under qualification testing by SSL leverages the substantial heritage of the SPT-100 being flown on Russian and US commercial satellites. The Jet Propulsion Laboratory is exploring the use of commercial EP systems, including the SPT-140, for deep space science missions, and initiated a program to evaluate the SPT-140 in the areas of low power operation and thruster operating life. A qualification model SPT-140 designated QM002 was evaluated for operation and plasma properties along channel centerline, from 4.5 kW to 0.8 kW. Additional testing was performed on a development model SPT-140 designated DM4 to evaluate operation with a Moog proportional flow control valve (PFCV). The PFCV was commanded by an SSL engineering model PPU-140 Power Processing Unit (PPU). Performance measurements on QM002 at 0.8 kW discharge power were 50 mN of thrust at a total specific impulse of 1250 s, a total thruster efficiency of 0.38, and discharge current oscillations of under 3% of the mean current. Steady-state operation at 0.8 kW was demonstrated during a 27 h firing. The SPT-140 DM4 was operated in closed-loop control of the discharge current with the PFCV and PPU over discharge power levels of 0.8-4.5 kW. QM002 and DM4 test data indicate that the SPT-140 design is a viable candidate for NASA missions requiring power throttling down to low thruster input power.

Garner, Charles E.↗
NASA NTRS1982

Simplification of power electronics for ion thruster neutralizers

A need exists for less complex and lower cost ion thruster systems. Design approaches and the demonstration of neutralizer power electronics for relaxed neutralizer keeper, tip heater, and vaporizer requirements are discussed. The neutralizer circuitry is operated from a 200 to 400 V bus and demonstrates an order of magnitude reduction in parts count. Furthermore, a new technique is described for regulating tip heater power and automatically switching over to provide keeper power with only four additional components. A new design to control the flow rate of the neutralizer with one integrated circuit is also presented.

Gruber, R. P.↗