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NASA NTRS1966

Thermal Control of Liquid H2/O2 Rocket Engines and Vehicle Impact

Cryogenic propulsion systems are subjected to various engine start environments during ground testing, first burn, and restart after orbital coast. These operational modes necessitate careful engine-stage integration, particularly in thermal control of propellant feed systems and engine subsystems. The wide range of operational conditions imposed on propellant feed systems include varying heat loads and acceleration forces for static test first burn and orbital restart. One method used to provide Net Positive Suctign,Head (NPSH) and maintain turbopump I is thermal control propellant circulation through the feed system prior to engine start. The attendant design and operational criteria are inter-related with stage and vehicle systems, including for example, auxiliary propulsion and propellant ullaging methods such as propellant tank propulsive venting. Thermally sensitive engine components require environmental control, protection, or special sequencing in order to assure repeatable engine start performance. An example is thrust chamber pre-chill to avoid pump stall during rapid start transient. Trends in engine development are toward expanding system flexibility to accommodate the wide variety of environments without special pre-start conditioning.

Winstead, Thomas W.↗

NASA NTRS1976

Comparison of vortex lattice predicted forces with wind tunnel experiments for the F-4E(CCV) airplane with a closely coupled canard

The F-4E (CCV) wind tunnel model with closely coupled canard control surfaces was analyzed by means of a version of a vortex lattice program that included the effects of nonlinear leading edge or side edge vortex lift on as many as four individual planforms. The results were compared with experimental data from wind tunnel tests of a 5% scale model tested at a Mach number M = 0.6. They indicated that a nonlinear vortex lift developed on the side edges due to tip vortices, but did not appear to develop on the leading edges within the range of angles of attack that were studied. Instead, substantial leading edge thrust was developed on the lifting surfaces. A configuration buildup illustrated the mutual interference between the wing and control surfaces. On the configuration studied, addition of the wing increased the loading on the canard, but the additional load on the canard due to adding the stabilator was small.

Gross, L. W.↗

NASA NTRS1959

Force-Test Investigation of the Stability and Control Characteristics of a 1/4-Scale Model of a Tilt-Wing Vertical-Take-Off-and-Landing Aircraft

A wind-tunnel investigation has been made to determine the aerodynamic characteristics of a 1/4-scale model of a tilt-wing vertical-take-off-and-landing aircraft. The model had two 3-blade single-rotation propellers with hinged (flapping) blades mounted on the wing, which could be tilted from an incidence of 4 deg for forward flight to 86 deg for hovering flight. The investigation included measurements of both the longitudinal and lateral stability and control characteristics in both the normal forward flight and the transition ranges. Tests in the forward-flight condition were made for several values of thrust coefficient, and tests in the transition condition were made at several values of wing incidence with the power varied to cover a range of flight conditions from forward-acceleration (or climb) conditions to deceleration (or descent) conditions The control effectiveness of the all-movable horizontal tail, the ailerons and the differential propeller pitch control was also determined. The data are presented without analysis.

Newsom, William A., Jr.↗

NASA NTRS2008

Preliminary Analysis for an Optimized Oil-Free Rotorcraft Engine Concept

Recent developments in gas foil bearing technology have led to numerous advanced high-speed rotating system concepts, many of which have become either commercial products or experimental test articles. Examples include Oil-Free microturbines, motors, generators and turbochargers. The driving forces for integrating gas foil bearings into these high-speed systems are the benefits promised by removing the oil lubrication system. Elimination of the oil system leads to reduced emissions, increased reliability, and decreased maintenance costs. Another benefit is reduced power plant weight. For rotorcraft applications, this would be a major advantage, as every pound removed from the propulsion system results in a payload benefit. Implementing foil gas bearings throughout a rotorcraft gas turbine engine is an important long-term goal that requires overcoming numerous technological hurdles. Adequate thrust bearing load capacity and potentially large gearbox applied radial loads are among them. However, by replacing the turbine end, or hot section, rolling element bearing with a gas foil bearing many of the above benefits can be realized. To this end, engine manufacturers are beginning to explore the possibilities of hot section gas foil bearings in propulsion engines. This paper presents a logical follow-on activity by analyzing a conceptual rotorcraft engine to determine the feasibility of a foil bearing supported core. Using a combination of rotordynamic analyses and a load capacity model, it is shown to be reasonable to consider a gas foil bearing core section.

Howard, Samuel A.↗

NASA NTRS2016

Electrospray Thruster Performance with Microfabricated Emitter Arrays and Indium Propellant

The feasibility of a microfabricated indium-fueled electrospray thruster with excellent performance was demonstrated. High efficiency electrospray thrusters with microfabricated components are under development for very compact, distributable propulsion systems that can be employed on both very small and large spacecraft with 10X improvement over SOA in mass, volume and specific impulse The critical components of this technology are the microfabricated emitter arrays and the capillary force driven indium feed system. Grey scale electron-beam lithography patterning and reactive ion dry etching provided the required micron-scale etch precision and uniformity across an array of 400 emitters in 1 cm2. Initial tests of single microfabricated silicon emitters demonstrated better performance than industry standard single emitters and the performance required to achieve 200 micronewtons when scaled up to a thruster with 400 emitters in 1 cm2. Arrays of emitters tested in a preliminary prototype Microfluidic Electrospray Propulsion (MEP) thruster assembly demonstrated stable performance at estimated thrust levels of 5, 50, and 120 µN at extraction voltages less than 4 kV. Current stability was within 0.15 % at 120 µN with only 1.5% of the emitter current collected by the extractor electrode. Post-test inspections revealed that more than 99% of the 400 emitters were electrospraying during the test. Specific impulse was estimated to be >3100 s from measurements of total charge and consumed indium mass at an emitter voltage of 1470 V. The results of this investigation suggest that microfabricated indium electrospray thruster technology is both feasible and capable of excellent performance as a highly compact microthruster.

Marrese-Reading, Colleen M.↗

NASA NTRS2003

Legacy and Emergence of Spaceport Technology Development at the Kennedy Space Center

Kennedy Space Center (KSC) has a long and successful legacy in the checkout and launch of missiles and space vehicles. These operations have become significantly more complex, and their evolution has driven the need for many technology developments. Unanticipated events have also underscored the need for a local, highly responsive technology development and testing capability. This evolution is briefly described, as well as the increasing level of technology capability at KSC. The importance of these technologies in achieving past national space goals suggests that the accomplishment of low-cost and reliable access to space will depend critically upon KSC's future success in developing spaceport technologies. This paper concludes with a description KSC's current organizational approach and major thrust areas in technology development. The first phase of our historical review focuses on the development and testing of field- deployable short- and intermediate-range ballistic missiles (1953 to 1958). These vehicles are later pressed into service as space launchers. The second phase involves the development of large space lift vehicles culminating in the Saturn V launches (1959 to 1975). The third phase addresses the development and operations of the partially reusable launch vehicle, Space Shuttle (1976 to 2000). In the current era, KSC is teaming with the U.S. Air Force (AF), industry, academia, and other partners to identify and develop Spaceport and Range Technologies necessary to achieve national space goals of lower-cost and higher-reliability space flight.

Starr, Stanley↗

NASA NTRS2007

Conceptual Design and Feasibility of Foil Bearings for Rotorcraft Engines: Hot Core Bearings

Recent developments in gas foil bearing technology have led to numerous advanced high-speed rotating system concepts, many of which have become either commercial products or experimental test articles. Examples include oil-free microturbines, motors, generators and turbochargers. The driving forces for integrating gas foil bearings into these high-speed systems are the benefits promised by removing the oil lubrication system. Elimination of the oil system leads to reduced emissions, increased reliability, and decreased maintenance costs. Another benefit is reduced power plant weight. For rotorcraft applications, this would be a major advantage, as every pound removed from the propulsion system results in a payload benefit.. Implementing foil gas bearings throughout a rotorcraft gas turbine engine is an important long-term goal that requires overcoming numerous technological hurdles. Adequate thrust bearing load capacity and potentially large gearbox applied radial loads are among them. However, by replacing the turbine end, or hot section, rolling element bearing with a gas foil bearing many of the above benefits can be realized. To this end, engine manufacturers are beginning to explore the possibilities of hot section gas foil bearings in propulsion engines. This overview presents a logical follow-on activity by analyzing a conceptual rotorcraft engine to determine the feasibility of a foil bearing supported core. Using a combination of rotordynamic analyses and a load capacity model, it is shown to be reasonable to consider a gas foil bearing core section. In addition, system level foil bearing testing capabilities at NASA Glenn Research Center are presented along with analysis work being conducted under NRA Cooperative Agreements.

Howard, Samuel A.↗

NASA NTRS2006

Holographic Solar Photon Thrusters

A document discusses a proposal to incorporate holographic optical elements into solar photon thrusters (SPTs). First suggested in 1990, SPTs would be systems of multiple reflective, emissive, and absorptive surfaces (solar sails) that would be attached to spacecraft orbiting the Earth to derive small propulsive forces from radiation pressures. An SPT according to the proposal would include, among other things, a main sail. One side of the sail would be highly emissive and would normally face away from the Earth. The other side would be reflective and would be covered by white-light holographic images that would alternately become reflective, transmissive, and absorptive with small changes in the viewing angle. When the spacecraft was at a favorable orbital position, the main sail would be oriented to reflect sunlight in a direction to maximize the solar thrust; when not in a favorable position, the main sail would be oriented to present a substantially absorptive/emissive aspect to minimize the solar drag. By turning the main sail slightly to alternate between the reflective and absorptive/ emissive extremes, one could achieve nearly a doubling or halving of the radiational momentum transfer and, hence, of the solar thrust.

Johnson, Les↗

NASA NTRS2011

The Dawn of Vesta Science

The Dawn mission is part of NASA's Discovery Program and has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total ?V of 11.3 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Fullpower thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional ?V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. An additional 210 hours of IPS thrusting was used to enter the first Vesta science orbit, called Survey orbit, on August 3, 2011 at an altitude of approximately 2,735 km. To date the IPS has been operated for 23,621 hours, consumed approximately 252 kg of xenon, and provided a delta-V of approximately 6.7 km/s. The IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. The only significant issue in over the almost four years of IPS operations in flight was the temporary failure of a valve driver board in the Digital Control and Interface Unit-1 (DCIU-1), resulting in a loss of thrust of approximately 29 hours. Thrusting operations resumed after switching to DCIU-2, and power cycling conducted after orbit capture restored DCIU-1 to full functionality. After about three weeks of Survey orbit operations the IPS will be used to transfer the spacecraft to the other planned science orbit altitudes. After approximately one year of science operations the IPS will be used for escape from Vesta and then for thrusting to Ceres with a planned arrival date at Ceres in February 2015. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations through the start of science operations at Vesta.

electric propulsion↗

NASA NTRS2018

Instability and Target Design in a Z-Pinch Driven Pulsed Fission / Fusion Engine

Advanced forms of propulsion are required to drastically improve upon the limits of existing technology and enable deep space exploration. Nuclear reactors, both fission and fusion, have great potential as systems with high energy density for spacecraft propulsion. There are a variety of nuclear propulsion concepts that have been proposed over the last several decades. Some of these rely on the confinement of plasmas by both magnetic and inertial means. When discussing the plasma confinement one must consider the stability of the system. The conditions of the plasma dictate the burn rate of the nuclear fuel; therefore, the plasma must maintain a minimum set of conditions in order to achieve an adequate burn. Instabilities limit plasma confinement through growing fluctuations in density and pressure. This results in turbulent mixing and loss of confinement. Thus it is desirable to find ways of managing the impact of instabilities that arise in order for a sufficient burn of the fuel such that surplus energy can be converted into a propulsive force. Note that the same instabilities plague the confinement of plasmas for terrestrial power generation. Over the last several years there has been research relating to the development of a Pulsed Fission-Fusion Engine (PuFF) conducted by the authors and others at their respective institutions. The propulsion system concept centers on the use of a z-pinch to compress a plasma in order to induce nuclear reactions. The plasma along with the additional energy released by the fission and fusion processes is expanded with a magnetic nozzle to produce thrust. The concept, if developed, is expected to have an improvement in specific impulse of several orders of magnitude in comparison with chemical propulsion systems. The z-pinch has been explored in the past as a method to confine a plasma via magnetic confinement fusion (MCF) for power generation. It suffers; however, from instabilities that so far have prevented break even conditions.

Taylor, Brian↗

NASA NTRS1997

Wings: A New Paradigm in Human-Centered Design

Many aircraft accidents/incidents investigations cite crew error as a causal factor (Boeing Commercial Airplane Group 1996). Human factors experts suggest that crew error has many underlying causes and should be the start of an accident investigation and not the end. One of those causes, the flight deck design, is correctable. If a flight deck design does not accommodate the human's unique abilities and deficits, crew error may simply be the manifestation of this mismatch. Pilots repeatedly report that they are "behind the aircraft" , i.e., they do not know what the automated aircraft is doing or how the aircraft is doing it until after the fact. Billings (1991) promotes the concept of "human-centered automation"; calling on designers to allocate appropriate control and information to the human. However, there is much ambiguity regarding what it mean's to be human-centered. What often are labeled as "human-centered designs" are actually designs where a human factors expert has been involved in the design process or designs where tests have shown that humans can operate them. While such designs may be excellent, they do not represent designs that are systematically produced according to some set of prescribed methods and procedures. This paper describes a design concept, called Wings, that offers a clearer definition for human-centered design. This new design concept is radically different from current design processes in that the design begins with the human and uses the human body as a metaphor for designing the aircraft. This is not because the human is the most important part of the aircraft (certainly the aircraft would be useless without lift and thrust), but because he is the least understood, the least programmable, and one of the more critical elements. The Wings design concept has three properties: a reversal in the design process, from aerodynamics-, structures-, and propulsion-centered to truly human-centered; a design metaphor that guides function allocation and control and display design; and a deliberate distinction between two fundamental functions of design, to complement and to interpret human performance. The complementary function extends the human's capabilities beyond his or her current limitations - this includes sensing, computation, memory, physical force, and human decision making styles and skills. The interpretive (or hermeneutic, Hollnagel 1991) function translates information, functionality, and commands between the human and the aircraft. The Wings design concept allows the human to remain aware of the aircraft through natural interpretation. It also affords great improvements in system performance by maximizing the human's natural abilities and complementing the human's skills in a natural way. This paper will discuss the Wings design concept by describing the reversal in the traditional design process, the function allocation strategy of Wings, and the functions of complementing and interpreting the human.

Schutte, Paul C.↗

NASA NTRS1993

Effects of discrete-source outgassing on motions of periodic comets and discontinuous orbital anomalies

Isoline maps of the Style II nongravitational parameters A1 and A2 of periodic comets are presented for a baseline model as functions of the parameters of the inertially fixed spin vector and the source's location-dependent thrust angle. The parameters A1 and A2 for a collection of sources are equal to the sum of their values for the individual contributors, thus depending on the source distribution. The model is consistent with even the most extreme among detected perturbations, including instances of rapid temporal changes in A2 and the discontinuous orbital anomalies, which are interpreted as products of a sudden redistribution of the transferred momentum due to episodic activation of short-lived sources. It is argued that the lifetimes of active regions vary considerably from case to case, and forced precession of the nucleus strongly influences the comet's activity pattern.

Sekanina, Zdenek↗

NASA NTRS1978

Flight effects on the aerodynamic and acoustic characteristics of inverted profile coannular nozzles

The effect of forward flight on the jet noise of coannular exhaust nozzles, suitable for Variable Stream Control Engines (VSCE), was investigated in a series of wind tunnel tests. The primary stream properties were maintained constant at 300 mps and 394 K. A total of 230 acoustic data points was obtained. Force measurement tests using an unheated air supply covered the same range of tunnel speeds and nozzle pressure ratios on each of the nozzle configurations. A total of 80 points was taken. The coannular nozzle OASPL and PNL noise reductions observed statically relative to synthesized values were basically retained under simulated flight conditions. The effect of fan to primary stream area ratio on flight effects was minor. At take-off speed, the peak jet noise for a VSCE was estimated to be over 6 PNdB lower than the static noise level. High static thrust coefficients were obtained for the basic coannular nozzles, with a decay of 0.75 percent at take-off speeds.

Kozlowski, H.↗

NASA NTRS1990

A computer module used to calculate the horizontal control surface size of a conceptual aircraft design

The creation of a computer module used to calculate the size of the horizontal control surfaces of a conceptual aircraft design is discussed. The control surface size is determined by first calculating the size needed to rotate the aircraft during takeoff, and, second, by determining if the calculated size is large enough to maintain stability of the aircraft throughout any specified mission. The tail size needed to rotate during takeoff is calculated from a summation of forces about the main landing gear of the aircraft. The stability of the aircraft is determined from a summation of forces about the center of gravity during different phases of the aircraft's flight. Included in the horizontal control surface analysis are: downwash effects on an aft tail, upwash effects on a forward canard, and effects due to flight in close proximity to the ground. Comparisons of production aircraft with numerical models show good accuracy for control surface sizing. A modified canard design verified the accuracy of the module for canard configurations. Added to this stability and control module is a subroutine that determines one of the three design variables, for a stable vectored thrust aircraft. These include forward thrust nozzle position, aft thrust nozzle angle, and forward thrust split.

Sandlin, Doral R.↗

NASA NTRS2015

In-Flight Operation of the Dawn Ion Propulsion System Through Survey Science Orbit at Ceres

The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total delta V of 11 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional delta V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in September 2012 and concluded with the start of the approach to Ceres phase on December 26, 2015, leading to orbit capture on March 6, 2015. Deterministic thrusting continued during approach to place the spacecraft in its first science orbit, called RC3, which was achieved on April 23, 2015. Following science operations at RC3 ion thrusting was resumed for twenty-five days leading to arrival to the next science orbit, called survey orbit, on June 3, 2015. The IPS will be used for all subsequent orbit transfers and trajectory correction maneuvers until completion of the primary mission in approximately June 2016. To date the IPS has been operated for over 46,774 hours, consumed approximately 393 kg of xenon, and provided a delta V of over 10.8 km/s to the spacecraft. The IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations through arrival at the second science orbit at Ceres.

Garner, Charles E.↗

NASA NTRS1998

Lift Fan Nozzle for Joint Strike Fighter Tested in NASA Lewis' Powered Lift Rig

Under a nonreimbursable space act agreement between the NASA Lewis Research Center and the Allison Advanced Development Company, Allison tested a lift fan nozzle in Lewis' Powered Lift Rig. This test was in support of the Joint Strike Fighter program (formerly the Joint Advanced Strike Technology) sponsored by the Department of Defense, which will develop and field an affordable, multirole, next-generation, strike fighter aircraft for the Navy, Air Force, Marine Corps, and foreign allies. Allison, along with Pratt & Whitney Company, is part of the Lockheed Martin Corporation team that is scheduled to build a concept demonstrator aircraft by fiscal year 2001. The test was initiated in April and successfully completed in mid-July of 1997. Allison supplied a one-third-scale model of the lift fan nozzle, and Lewis provided the facility and the necessary support team. Various configurations, including pitching vectored angles ranging from 15deg forward to 60deg backward, were tested over a range of nozzle pressure ratios. Nozzle flow rates, thrust, and static pressures were measured for each of the configurations. Results from the test met the design requirements for the Joint Strike Fighter program and were in agreement with Allison's internal computational fluid dynamics (CFD) analyses. Data obtained from this test will also be used in the full-scale design of the lift fan system.

Lam, David W.↗

NASA NTRSDate not supplied

Benefit and Critical Factors for the Performance of the Boundary Layer Ingesting Propulsion

The benefit of the boundary layer ingestion (BLI) is described in the perspective of the propulsion and engine development. A power saving map of the BLI engines is derived based on the correlation of the shape factor of the ingested boundary layer and the propulsive efficiency. The ratio of the mass flow rate between BLI and non-BLI propulsors is introduced to quantify the power saving of the BLI engine relative to a corresponding clean inlet flow engine that generates the same amount of net thrust. The wake recovery factor from the jet flow is applied to find the optimum sizing of the engine for the given design requirement. The effects of the fan pressure ratio of the propulsor toward the power saving are also investigated to seek out the feasible range of the BLI propulsor design. The derived correlation is validated with CFD analyses. Simulation models using variously sized engines relative to an influencing body are developed so that a numerical experiment is carried out with the various shape factors. The effect of the propulsion airframe integration toward the aerodynamic forces on the fuselage and nacelle is investigated via parametric study of the engine sizing and fan pressure ratio for the wake ingestion models. The propulsor efficiency is quantified toward the saving in the actual shaft power and correlated with the efficiency penalty of the BLI propulsor. The correlation is also validated via comparison of the turbo-machinery CFD results for BLI2DTF and conventional transonic fan stages.

Boundary Layer Ingestion↗

NASA NTRSDate not supplied

Advancements in Force and Angle Measurement by the NFMTC

The National Force Measurement Technology Capability (NFMTC) is a nationwide partnership established in 2008 and sponsored by NASA’s Aeronautics Evaluation and Test Capabilities (AETC) project to maintain and further develop force and angle measurement capabilities primarily for NASA’s ground test facilities. One of the NFMTC’s primary objectives is to develop new force and angle measurement capabilities to meet the needs of researchers testing in NASA’s ground test facilities. This work will highlight several new capabilities that are either at or near operational readiness. A new six-component flow-through balance was designed and fabricated to support an upcoming retropropulsion test campaign at the NASA Langley Unitary Plan Wind Tunnel. The balance employs membranes in place of bellows, which have historically been used, and passes the flow through load bearing parts of the balance. Flow through the balance is optimized via the addition of flow guides, which are manufactured as separate parts, and attached to the balance. Calibration of this new flow-through balance demonstrates that it can suitably characterize the load and pressure ranges it was designed for, which provides aero researchers a new tool to directly measure retropropulsion performance of powered descent vehicles. Balances used in NASA’s ground test facilities are typically either single-piece balances or floating-frame balances. Historically, these two types of balances have generally used different front-end (model) attachments. Single-piece balances generally have a cylindrical front end that is forward of any measuring sections and at a reduced diameter relative to the balance outer diameter. Floating-frame balances generally have front end fits that extend over one or more measuring sections and that engage at the balance outer diameter. Recently, the NFMTC fabricated a single-piece balance with a front end fit that is compatible with the hardware used to calibrate floating-frame balances. Calibration results evaluating this front end fit will be presented. From a recent survey of angle measurement needs, the development of smaller angle measurement packages was identified as one area of improvement. Micro-electromechanical systems (MEMS) are continually being refined to provide high performance accelerometers that are a fraction of the volume of the quartz flexure based accelerometers that are standard in wind tunnels. A survey of commercial off-the-shelf MEMS accelerometers was conducted, and it identified the Safran Colibrys MS lines to be of adequate performance. Single-axis test boards and tri-axis rigid-flex production boards were designed and fabricated with a volume of approximately .18 cubic inches. The boards were designed to provide all the necessary conditioning for the MEMS accelerometers at levels to meet or exceed all requirements for the MS line sensors. Preliminary testing of these assembled sensor packages has demonstrated acceptable performance for wind-tunnel angle measurement. Several new rotating balances are at various stages of completion for NASA Glenn Research Center’s fan drive rigs. These balances provide direct measurement of thrust and torque of a driven fan or fan section to provide an estimate of fan efficiency and net thrust. Three two-component balances have been recently completed, and one six-component balance is in active development. A custom rotating telemetry system has been developed to support testing with these balances, and the telemetry system is integrated into the forward end of each one. The telemetry system is designed to remove mechanical sliprings by using an optical ethernet-based slipring for data transmission.

Force measurement↗