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Ultra-Efficient Engine Technology (UEET) Proof of Concept Compressor, Advanced Compressor Casing Treatment testing, First Research Configuration (concentric grooves)

Ultra-Efficient Engine Technology (UEET), Proof of Concept Compressor, Advanced Compressor Casing Treatment testing; close up - throttle valve -wide open

Ultra-Efficient Engine Technology (UEET), Proof of Concept Compressor, Advanced Compressor Casing Treatment testing; close up - throttle valve -wide open

One of the most important properties of aviation fuels for spark-ignition engines is their knock rating. The CFR engine tests of fuels of 87 octane and above does not always correspond entirely to the actual behavior of these fuels in the airplane engine. A method is therefore developed which, in contrast to the octane number determination, permits a testing of the fuel under various temperatures and fuel mixture conditions. The following reference fuels were employed: 1) Primary fuels; isooctane and n-heptane; 2) Secondary fuels; pure benzene and synthetic benzine.

The Augmented Deflector Exhaust Nozzle (ADEN) exhaust system was tested in the PSL-3 altitude chamber at the NASA Lewis Research Center in order to evaluate aerodynamic performance, cooling-system effectiveness, and mechanical operation at flight-type conditions. The ADEN, a flight-weight, two-dimensional, thrust-vectoring nozzle, was successfully tested on the F404 engine using a remote engine control system for automatic or manual setting of the throat-area control and available fan air for the nozzle internal cooling system. Throughout the test, the ADEN performed with no adverse effects on the engine or augmentor operation.

The performance of a turbocharged rotary engine at power levels above 75 kW (100 hp) was studied. A twin rotor turbocharged Mazda engine was tested at speeds of 3000 to 6000 rpm and boost pressures to 7 psi. The NASA developed combustion diagnostic instrumentation was used to quantify indicated and pumping mean effect pressures, peak pressure, and face to face variability on a cycle by cycle basis. Results of this testing showed that a 5900 rpm a 36 percent increase in power was obtained by operating the engine in the turbocharged configuration. When operating with lean carburetor jets at 105 hp (78.3 kW) and 4000 rpm, a brake specific fuel consumption of 0.45 lbm/lb-hr was measured.

Ultra-Efficient Engine Technology (UEET), Proof of Concept Compressor, Advanced Compressor Casing Treatment testing; bearing housing and shaft during build-up

Ultra-Efficient Engine Technology (UEET) Proof of Concept Compressor, Advanced Compressor Casing Treatment testing, First Research Configuration (concentric grooves)

Jet engine being set up for testing

Jet engine being set up for testing

Optical fire detector systems for aircraft engine compartments provided by fire vendors were evaluated in the Aircraft Engine Compartment Fire Test Simulator (AENFTS). Their ability to sense JP-4 fires of two sizes and provide fire warnings in a variety of simulated flight environments was monitored along with their freedom from false fire warnings. One systems was identified which provided fire warnings for every test fire and was completely free of false alarms.

Ultra-Efficient Engine Technology (UEET), Proof of Concept Compressor, Advanced Compressor Casing Treatment testing; bearing housing and shaft during build-up

Many dynamic systems operate in select operating regions, each exhibiting characteristic modes of behavior. It is traditional to employ standard adjustable gain proportional-integral-derivative (PID) loops in such systems where no apriori model information is available. However, for controlling inlet pressure for rocket engine testing, problems in fine tuning, disturbance accommodation, and control gains for new profile operating regions (for research and development) are typically encountered. Because of the capability of capturing I/O peculiarities, using NETS, a back propagation trained neural network is specified. For select operating regions, the neural network controller is simulated to be as robust as the PID controller. For a comparative analysis, the higher order moment neural array (HOMNA) method is used to specify a second neural controller by extracting critical exemplars from the I/O data set. Furthermore, using the critical exemplars from the HOMNA method, a third neural controller is developed using NETS back propagation algorithm. All controllers are benchmarked against each other.

Metallic contaminant was found in the liquid oxygen (LOX) pre-valve screen of the shuttle main engine propulsion system on two orbiter vehicles. To investigate the potential for an ignition, NASA Johnson Space Center White Sands Test Facility performed (modified) rotating friction ignition testing in LOX. This testing simulated a contaminant particle in the low-pressure oxygen turbo pump (LPOTP) and the high-pressure oxygen turbo pump (HPOTP) of the shuttle main propulsion system. Monel(R) K-500 and Inconel(R) 718 samples represented the LPOTP and HPOTP materials. Aluminum foil tape and titanium foil represented the contaminant particles. In both the Monel(R) and Inconel(R) material configurations, the aluminum foil tape samples did not ignite after 30 s of rubbing. In contrast, all of the titanium foil samples ignited regardless of the rubbing duration or material configuration. However, the titanium foil ignitions did not propagate to the Monel and Inconel materials.

Based on the Apollo-era J-2 that powered the second and third stages of the Saturn V, the current J-2X is the liquid hydrogen and oxygen high-altitude rocket engine in development for both the Ares I Upper Stage and Ares V Earth Departure Stage. During my summer 2009 internship, J-2X was at a stage in its design maturity where verification testing needed to be considered for the benefit of adequate test facility preparation. My task was to focus on gimbal requirements and gimbal related hot-fire test plans. Facility capabilities were also of interest, specifically for hot-fire testing slated to occur at test stands A-1, A-2, and A-3 at Stennis Space Center(SSC) in Bay St. Louis, Mississippi. Gimbal requirements and stage interface conditions were investigated by applying a top-to-bottom systems engineering approach, which involved system level requirements, engine level requirements from both government and engine contractor perspectives, component level requirements, and the J-2X to Upper Stage and Earth Departure Stage interface control documents. Previous hydrogen and oxygen liquid rocket engine gimbal verification methods were researched for a glimpse at lessons learned. Discussion among the J-2X community affected by gimballing was organized to obtain input relative to proper verification of their respective component. Implementing suggestions such as gimbal pattern, angulated dwell time, altitude testing options, power level, and feed line orientation, I was able to match tests to test stands in the A Complex at SSC. Potential test capability gaps and risks were identified and pursued. The culmination of all these efforts was to coordinate with SSC to define additional facility requirements for both the A-3 altitude test stand that is currently under construction and the A-1 sea level test stand which is being renovated

Rotating detonation engines (RDEs) have the potential for greater efficiencies over conventional engines by utilizing pressure gain combustion. A new modular RDE (6 in diameter) was developed and successfully run on hydrogen and standard air. The RDE allows for variation of injection scheme and detonation channel widths. Tests provided the operational space of the new RDE as well as characterized detonation unsteadiness. It was found that a smaller equivalence ratio than previous was required to obtain continuous detonations. Also discovered was VCJ was reached in the RDE, but not sustained.

Approved For ReF se 2000/08/29 : CIA-RDP79RO0961AOO0800040014-1 . SECRET 25X1 C Possible Rocket Engine Static Testing: similar value. The source was queried first on noises he reported on in an earlier interrogation. Source lived at from Spring 1948 to July 1950, 25X1C from July 1950 to February 1951, and - from February 1951 25X1 C to September 1955. Noises of extreme intensity, sufficient to rattle their windows first occurred in late 1,048 or early 1949. These occurred apnroxi_- mately 15 to 20 times in two and one half years at at no set time 25X1 C of day. The sound was of a rumbling nature with a high hiss background. The azimuth of the sound was slightly north of due east. Nothing was visible to the eye from their location. The duration of the noise was approximately thirty seconds, followed by a several second absence and then a resumption of about thirty seconds. The intensity of noises observed always were of a 25X1 C Although source had a short stay at identical noises but of lower intensity were noted. The azimuth of the sound was approximated at about forty-five degree east of north. 25X1 C Identical noises were noted at their - location of intensity level 25X1 C similar to The azimuth of sound emission was approximated north to just barely east of north. Although not certain, he believes that the noises were less frequent by 1953/54 and may have ceased by the summer of 1954. Source believed that noises were attributable to rockets. He based this on a former experience in the 1930s when he observed tests of Fritz von Opel's rocket car experiments. Source pointed out that he was in Moscow six days a week and therefore many tests (noises) may have occurred in his absence. SECRET Approved For Release 2000/08/29 : CIA-RDP79ROO961AO00800040014-1

This report details the use of infrared imaging for temperature measurement and thermal pattern determination during simulated altitude engine testing in the NASA Lewis Propulsion Systems Laboratory. Three identical argon-cooled imaging systems were installed in the facility exhaust collector behind sapphire windows to look at engine internal surfaces. The report describes the components of each system, presents the specifics of the complicated installation, and explains the operation of the systems during engine testing. During the program, several problems emerged, such as argon contamination system, component overheating, cracked sapphire windows, and other unexplained effects. This report includes a summary of the difficulties as well as the solutions developed. The systems performed well, considering they were in an unusually harsh exhaust environment. Both video and digital data were recorded, and the information provided valuable material for the engineers and designers to quickly make any necessary design changes to the engine hardware cooling system. The knowledge and experience gained during this program greatly simplified the installation and use of the systems during later test programs in the facility. The infrared imaging systems have significantly enhanced the measurement capabilities of the facility, and have become an outstanding and versatile testing resource in the Propulsion Systems Laboratory.

Jet engine being set up for testing

This document reports an experimental study of the effectiveness of dry chemical in extinguishing and delaying reignition of fires resulting from hydrocarbon fuel leaking onto heated surfaces such as can occur in jet engine nacelles. The commercial fire extinguishant dry chemicals tried were, for example, sodium and potassium bicarbonate, carbonate, chloride, and carbamate (Monnex) but we have also tested other metal-halogen and metal-hydroxycarbonate compounds prepared in our own laboratories. Given in this report are: synthetic and preparative procedures for new materials developed; a new concept of fire-control by dry chemical agents; descriptions of experimental assemblages to test dry chemical fire extinguishant efficiencies in controlling fuel fires initiated by hot surfaces; comparative testing data for more than 25 chemical systems in a static assemblage with no airflow across the heated surface, and similar comparative data for more than 10 compounds in a dynamic system with airflows up to 350 ft/sec; and recommendations for future work with one system that fulfills all requirements delineated by the sponsoring agency, and which has been tested in both the Static and Dynamic assemblages with both methodologies confirming it as the most effective system by comparison with the other materials tested.

NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder-bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. NASA's efforts include a 4K lbf thrust liquid oxygen/methane (LOX/CH4) combustion chamber and subscale thrust chambers for 1.2K lbf LOX/hydrogen (H2) applications that have been designed and fabricated with SLM GRCop-84. The same technologies for these lower thrust applications are being applied to 25-35K lbf main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

No abstract available

The fan case in a jet engine is a heavy structure because of its size and because of the requirement that it contain a blade released during engine operation. Composite materials offer the potential for reducing the weight of the case. Efficient design, test, and analysis methods are needed to efficiently evaluate the large number of potential composite materials and design concepts. The type of damage expected in a composite case under blade-out conditions was evaluated using a subscale test in which a glass/epoxy composite half-ring target was impacted with a wedge-shaped titanium projectile. Fiber shearing occurred near points of contact between the projectile and target. Delamination and tearing occurred on a larger scale. These damage modes were reproduced in a simpler test in which flat glass/epoxy composites were impacted with a blunt cylindrical projectile. A surface layer of ceramic eliminated fiber shear fracture but did not reduce delamination. Tests on 3D woven carbon/epoxy composites indicated that transverse reinforcement is effective in reducing delamination. A 91 cm (36 in.) diameter full-ring sub-component was proposed for larger scale testing of these and other composite concepts. Explicit, transient, finite element analyses indicated that a full-ring test is needed to simulate complete impact dynamics, but simpler tests using smaller ring sections are adequate when evaluation of initial impact damage is the primary concern.

An uncertainty analysis was conducted to determine the bias and precision errors and total uncertainty of measured turbojet engine performance parameters. The engine tests were conducted as part of the Uniform Engine Test Program which was sponsored by the Advisory Group for Aerospace Research and Development (AGARD). With the same engines, support hardware, and instrumentation, performance parameters were measured twice, once during tests conducted in test cell number 3 and again during tests conducted in test cell number 4 of the NASA Lewis Propulsion Systems Laboratory. The analysis covers 15 engine parameters, including engine inlet airflow, engine net thrust, and engine specific fuel consumption measured at high rotor speed of 8875 rpm. Measurements were taken at three flight conditions defined by the following engine inlet pressure, engine inlet total temperature, and engine ram ratio: (1) 82.7 kPa, 288 K, 1.0, (2) 82.7 kPa, 288 K, 1.3, and (3) 20.7 kPa, 288 K, 1.3. In terms of bias, precision, and uncertainty magnitudes, there were no differences between most measurements made in test cells number 3 and 4. The magnitude of the errors increased for both test cells as engine pressure level decreased. Also, the level of the bias error was two to three times larger than that of the precision error.

Five firings of the J-2S rocket engine were conducted in Rocket Development Test Cell (J-4) on July 17, and 29, 1969. These firings were accomplished at pressure altitudes ranging from 85,000 to 101,000 ft at engine start. The primary objectives of these test periods were to (1) determine if main-stage conditions which existed during sea-level testing of engine S/N J-113 would result in similar abnormal oxidizer dome vibrations in the 4400- to 4700- Hz frequency range during altitude testing of engine J-112-E, (2) evaluate high thrust idle-mode operation with a simulated full-face oxidizer flow injector configuration, and (3) document effects of closing the thrust chamber bypass valve during high thrust idle-mode operation. Altitude testing did not result in abnormal (greater than 100 grams) oxidizer dome vibration in the 4400- to 4700- Hz range during test period 04. The thrust chamber bypass valve closing resulted in a 65 F increase in fuel injection temperature; however, stabilized high thrust idle-mode operation was not attained during test period 05.

The objective of the demonstration test program was to confirm the validity of the new component designs by engine tests at both nominal and extreme Titan IIIM operating conditions without unnecessarily repeating the Titan II and Titan III qualification tests. The test program also served as a convenient vehicle for testing related items, not intrinsic to the engine system, such as the Martin POGO accumulators, gimbal actuators, flight instrumentation, and ground equipment. The planned program consisted of 30 tests, 15 on each of two engines.

Four firings, and one engine start to expiration of the ignition phase timer, of the Rocketdyne J-2 rocket engine were conducted in Test Cell J- 4. The firings were accomplished during test period J4-1801-02 at pressure altitudes from 97,000 to 106,000 ft at engine start. The objectives of the test were to evaluate S-IVB/S-V start condition effects on (1) gas generator and augmented spark igniter chamber ignition characteristics and (2) fuel pump stall characteristics during start tank blowdown for J-2 engine J-2052. Engine thermal conditions predicted for the J-2 engine flight configuration first burn, and restarts after one and two orbits, were simulated.

Four firings of the J-2 rocket engine were conducted in Test Cell J-4 of the Large Rocket Facility. The firings were accomplished during test period J4-1801-08 at pressure altitudes ranging from 92,500 to 107,000 ft at engine start. The objectives of the test were to evaluate S-IVB/S-V start condition effects on (1) engine start transients, (2) gas generator outlet temperature; (3) augmented spark igniter operation, and (4) fuel pump high level stall margin for J-2 engine S/N J-2052. The accumulated firing duration was 70.32 sec. Satisfactory engine operation was obtained.

Nine firings of the J-2 rocket engine are reported. The firings were accomplished at pressure altitudes ranging from 95,000 to 112,000 ft at engine start. The objectives of these tests were to investigate S-II/S-V start condition effects on engine start, gas generator ignition characteristics, and fuel pump operation. Engine thermal condition extremes for S-II flight configuration were duplicated. Satisfactory engine operation was obtained. Accumulated firing duration was 111.6 sec.

Ten idle-mode firings of the J-2S rocket engine S/N J-112 were conducted during test periods J4-1001-08 through J4-1001-10, and J4-1001-12 through J4-1001-14 between September 2 and October 15, 1969, in Rocket Development Test Cell (J-4) of the Engine Test Facility. The primary objectives of these firings were to determine engine idle-mode operating characteristics at various engine mixture ratios utilizing a new noncompartmented injector. Engine performance, as indicated by characteristic velocity, was 120 percent higher than experienced with previous compartmented injector designs at nominal pump inlet conditions. Total accumulated engine idle-mode operation was 1408.2 sec.

Ten idle-mode firings of the J-2S rocket engine S/N J-112 were conducted during test periods J4-1001-08 through J4-1001-10, and J4-1001-12 through J4-1001-14 between September 2 and October 15, 1969, in Rocket Development Test Cell (J-4) of the Engine Test Facility. The primary objectives of these firings were to determine engine idle-mode operating characteristics at various engine mixture ratios utilizing a new noncompartmented injector. Engine performance, as indicated by characteristic velocity, was 120 percent higher than experienced with previous compartmented injector designs at nominal pump inlet conditions. Total accumulated engine idle-mode operation was 1408.2 sec.

Performance profiling and testing is one of the interesting topics in the big data management and Cloud Computing. In testing, we use test cases composed to different type of queries to evaluate the performance aspects of the information retrieval system for large scale information collection. This test scenarioperforms the evaluation ofretrieval accuracy for all kind of ambiguity and non factoid queries with result set as Training data. This stands difficult to evaluate the retrieval method in order to schedule or optimize the Recommendation and prediction technique of the IR method to the Real time queries. The Queries is considered as requirement specification which has to supply to search engine or web information provider applications for information or web page retrieval. In this paper, we propose a novel technique named as “Test Retrieval Framework“a performance profiling and testing of the web search engines on the information retrieved towards non factoid queries. In this technique, we apply expectation maximization algorithm as an iterative method to find maximum likelihood estimate. We discuss on the important aspects in this work based on Recommendation models integrating domain and web usage, Query optimization for navigational and Transactional queries, Query Result records. The Experimental results demonstrates the proposed technique outperforms of state of arts approaches in terms of set based measures like Precision, Recall and F measure and rank based measures like Mean Average Precision and Cumulative Gain. 

Ultra-Efficient Engine Technology (UEET), Proof of Concept Compressor, Advanced Compressor Casing Treatment testing; bearing housing installed in collector with instrumentation.

Environmental legislation is being drafted by the EPA to reduce small particles in turbine engine exhaust. Six of the most promising fuel additives were chosen for a comparative evaluation as to the material compatibility with the non-metallic materials present on aircraft. These additives were added to JP-8+100 fuel prior to testing. In general, all results showed no significant concerns with the materials tested.

A new procedure for calibrating on-board gauges on the blade assemblies of turbomachines is proposed in this paper. The calibration of an engine gauge, determined a priori during a component dynamic test, makes it possible to establish the levels of dynamic stress on the entire instrumented airfoil (based on the readings output by the gauge during the engine test). The method proposed is based on the use of holographic dynamic measurements enabling the qualification of a finite element model in relation to the component test. The calibration relations are then determined based on the qualified model. The procedure proposed enables an increase in the reliability of the calibration relations by making it possible to take into account in the model the real conditions of the engine test (thermal, centrifugal and aerodynamic loads, and the shapes of dynamic modes of the complete blade assembly). Similarly, it enables a drastic reduction in the costs relating to the component tests required to establish these relations.

Measurable quantities of NOx, CO and small particulates are produced and are emitted into the atmosphere during the testing of aircraft engines in jet engine test cells (JETCs). These emissions have been and are a concern to the Air Force and to others who test aircraft engines. The large quantities of exhaust gases that are generated, the wide range of testing conditions that are normally employed, and the sensitivity of engines to back pressures make control difficult and the use of conventional control technologies impractical. A need exists for a simple, low-cost method to control the emissions. In a Phase I SBIR project, Sorbent Technologies Corporation (Sorbtech) explored the ability of vermiculite to reduce or capture contaminants in exhaust gas streams. During the Phase II SBIR project described in this report, Sorbtech investigated how vermiculite might be employed in a commercial system to control emissions from JETCs and how chemical additions to vermiculite might enhance its NOx-removal abilities. The objectives of the Phase II project were to develop and to demonstrate a suitable filter design involving vermiculite that will control NOx, CO, and small-particulate emissions during jet-engine testing. This report traces the development of a new control technology through the laboratory, bench-scale, slipstream, and prototype stages of testing and demonstration. The result of this work is a simple filter design that was recently evaluated at Tyndall AFB. The new filter design consists of two thin panel beds that capture contaminants as gas passes through them at the end of the exhaust-gas chimney. The filters consist of a primary bed of vermiculite/MgO and a secondary bed of virgin vermiculite or activated carbon. The secondary bed is placed in front of the primary bed. The filters were found to remove 40 to 83 percent of the NOx, more than 50 percent of the particulates by mass, and significant amounts of CO.

It is recommended that the testing completed with Engine 828 be accepted as evidence that the F107-WR-400 engine meets or exceeds the requirements for environmental vibration exposure and mission simulation testing as set forth in the PID Specification, 24235WR-9501A, and the Qualification Test Plan, CMEP 91-4043G, Report No. 78-145-8.

Engineers at the Marshall Space Flight Center install the F-1 engines on the S-IC stage thrust structure at the S-IC static test stand. Engines are installed on the stage after it has been placed in the test stand. Five F-1 engines, each weighing 10 tons, gave the booster a total thrust of 7,500,000 pounds, roughly equivalent to 160 million horsepower.

The fringe type laser anemometer is analyzed using the Cramer-Rao bound for the variance of the estimate of the Doppler frequency as a figure of merit. Mie scattering theory is used to calculate the Doppler signal wherein both the amplitude and phase of the scattered light are taken into account. The noise from wall scatter is calculated using the wall bidirectional reflectivity and the irradiance of the incident beams. A procedure is described to determine the optimum aperture mask for the probe volume located a given distance from a wall. The expected performance of counter type processors is also discussed in relation to the Cramer-Rao bound. Numerical examples are presented for a coaxial backscatter anemometer.

The propulsion and Energetics Panel, Working Group 15, of the Advisory Group for Aerospace Research and Development (AGARD) is sponsoring a Uniform Engine Testing Program (UETP). In this program, two jet engines were tested under identical conditions in certain NATO altitude and ground-level facilities as a means of correlating these facilities. With this second entry, NASA documented engine deterioration that may have occurred since inception of the UETP. Additionally, NASA investigated anomalies discovered during review of data from the five facilities which had participated in the program between the two NASA entries.

The fatigue type testing performed on instrumentation rakes and probes intended for use in the air flow passages of jet engines during full scale engine tests is outlined. A discussion of each type of test performed, the results that may be derived and means of inspection is included.

The design of a subscale jet engine driven ejector/diffuser system is examined. Analytical results and preliminary design drawings and plans are included. Previously developed performance prediction techniques are verified. A safety analysis is performed to determine the mechanism for detonation suppression.

Two builds of an F100 engine model derivative (EMD) engine were evaluated for improvements in engine components and digital electronic engine control (DEEC) logic. Two DEEC flight logics were verified throughout the flight envelope in support of flight clearance for the F100 engine model derivative program (EMPD). A nozzle instability and a faster augmentor transient capability was investigated in support of the F-15 DEEC flight program. Off schedule coupled system mode fan flutter, DEEC nose-boom pressure correlation, DEEC station six pressure comparison, and a new fan inlet variable vane (CIVV) schedule are identified.

Since 1984, an effort has been underway at Rocketdyne, manufacturer of the Space Shuttle Main Engine (SSME), to automate much of the analysis procedure conducted after engine test firings. Previously published articles at national and international conferences have contained the context of and justification for this effort. Here, progress is reported in building the full system, including the extensions of integrating large databases with the system, known as Scotty. Inductive knowledge acquisition has proven itself to be a key factor in the success of Scotty. The combination of a powerful inductive expert system building tool (ExTran), a relational data base management system (Reliance), and software engineering principles and Computer-Assisted Software Engineering (CASE) tools makes for a practical, useful and state-of-the-art application of an expert system.

this is an old vid, check this to see the fixed textures http://www.youtube.com/watch?v=Tpotj2ig6iI

Off-the-shelf jet propulsion in the 50 - 500 lb thrust class sparse. A true twin-spool turbofan in this range does not exist. Adapting an off-the-shelf turboshaft engine is feasible. However the approx.10 Hp SPT5 can t quite make 50 lbs. of thrust. Packaging and integration is challenging, especially the exhaust. Building on our engine using a 25 Hp turboshaft seems promising if the engine becomes available. Test techniques used, though low cost, adequate for the purpose.

Virgil Allen testing Dynafog Model 70 fog machine used for pest control. Machine has jet engine and disseminates only oil-based materials. Shelton Ranger District, Olympic National Forest, Washington. Photo by: Ken H. Wright Date: May 4, 1962 Credit: USDA Forest Service, Pacific Northwest Region, State and Private Forestry, Forest Health Protection. Collection: Portland Station Collection; La Grande, Oregon. Image: PS-2677 To learn more about this photo collection see: Wickman, B.E., Torgersen, T.R. and Furniss, M.M. 2002. Photographic images and history of forest insect investigations on the Pacific Slope, 1903-1953. Part 2. Oregon and Washington. American Entomologist, 48(3), p. 178-185. For additional historic forest entomology photos, stories, and resources see the Western Forest Insect Work Conference site: wfiwc.org/content/history-and-resources Image provided by USDA Forest Service, Pacific Northwest Region, State and Private Forestry, Forest Health Protection: www.fs.usda.gov/main/r6/forest-grasslandhealth

The design of a subscale jet engine driven ejector/diffuser system is examined. Analytical results and preliminary design drawings and plans are included. Previously developed performance prediction techniques are verified. A safety analysis is performed to determine the mechanism for detonation suppression.

The National Aeronautics and Space Administration (NASA) conducted a full scale ice crystal icing turbofan engine test using an obsolete Allied Signal ALF502-R5 engine in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The test article used was the exact engine that experienced a loss of power event after the ingestion of ice crystals while operating at high altitude during a 1997 Honeywell flight test campaign investigating the turbofan engine ice crystal icing phenomena. The test plan included test points conducted at the known flight test campaign field event pressure altitude and at various pressure altitudes ranging from low to high throughout the engine operating envelope. The test article experienced a loss of power event at each of the altitudes tested. For each pressure altitude test point conducted the ambient static temperature was predicted using a NASA engine icing risk computer model for the given ambient static pressure while maintaining the engine speed.

Flight tests were recently completed to demonstrate an inlet-distortion-tolerant engine control system. These flight tests were part of NASA's High Stability Engine Control (HISTEC) program. The objective of the HISTEC program was to design, develop, and flight demonstrate an advanced integrated engine control system that uses measurement-based, real-time estimates of inlet airflow distortion to enhance engine stability. With improved stability and tolerance of inlet airflow distortion, future engine designs may benefit from a reduction in design stall-margin requirements and enhanced reliability, with a corresponding increase in performance and decrease in fuel consumption. This paper describes the HISTEC methodology, presents an aircraft test bed description (including HISTEC-specific modifications) and verification and validation ground tests. Additionally, flight test safety considerations, test plan and technique design and approach, and flight operations are addressed. Some illustrative results are presented to demonstrate the type of analysis and results produced from the flight test program.