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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Ultraviolet photographic pyrometer investigates the role of carbon as a thermal radiator and determines the geometry, location, and progress of afterburning phenomena in the exhaust plume of rocket engines using liquid oxygen/RP-1 as propellant.
Wind tunnel for large component testing uses a jet engine with afterburner to provide high temperatures /1200 degrees to 2000 degrees F/ and controlled high velocity gas. This economical wind tunnel can accommodate parts ten feet by ten feet or larger, and is a useful technique for qualitative information.
Afterburning turbojet, nonafterburning turbojet, and duct burning turbofan designs for supersonic transports with and without takeoff noise constraints
Heat balance equation for predicting heat transfer of ejector exhaust nozzle at afterburning phase
Wall temperature and heat transfer characteristics of air cooled plug nozzle system for afterburning turbojet engine
Afterburner-equipped jet engine nacelle with slotted configuration afterbody
A calculational procedure is presented to analyze the heat-transfer and fluid-flow characteristics of a convectively air -cooled plug-nozzle operating on an afterburning turbojet engine. Anderson's method was used to predict hot-gas static pressures in the supersonic stream with fully expanded flow (high nozzle-pressure ratios); the results were excellent. For low nozzle-pressure ratios, the flow was assumed to expand one-dimensionally and isentropically to the plug back pressure. Wall temperatures predicted using this latter pressure distribution agreed well with the wall temperatures predicted using the measured hot-gas pressures (maximum deviation was about 30 K (54 deg R)). Either an in tegral boundary-layer technique or a simple pipe-flow equation may be used to calculate convective heat transfer from the hot gas to the wall. The simple pipeflow equation results in the prediction of slightly higher wall temperatures than does the integral technique. Experimental wall temperatures were generally in good agreement with the two predicted wall temperature distributions. Excellent agreement was noted b etween measured and predicted coolant static-pressure distributions. The plug-coolant temperature rise was generally overpredicted by about 22.2 K (40 deg R); possible explanations are offered. Although an an alysis of the struts, which support the plug, was purposely kept simple, reasonable results were obtained. Potential flow over an ellipse was used to calculate hot-gas static pressure; the results were satisfactory.
Twin-jet afterbody models were investigated by using two balances to measure the thrust-minus-total drag and the afterbody drag, separately, at static conditions and at Mach numbers up to 2.2 for an angle of attack of 0 deg. Hinged-flap convergent-divergent nozzles were tested at subsonic-cruise- and maximum-afterburning-power settings with a high-pressure air system used to provide jet-total-pressure ratios up to 20. Two nozzle lateral spacings were studied, using afterbodies with similar interfairing shapes but with different longitudinal cross-sectional area distributions. Alternate, blunter, interfairings with different shapes for the two spacings, which produced afterbodies having identical cross-sectional area progressions corresponding to an axisymmetric minimum wave-drag configuration, were also tested. The results indicate that the wide-spaced configurations improved the flow field around the nozzles, thereby reducing drag on the cruise nozzles; however, the increased surface and projected cross-sectional areas caused an increase in afterbody drag. Except for a slight advantage with cruise nozzles at subsonic speeds, the wide-spaced configurations had the higher total drag at all other test conditions.
Estimates of the prospects for quiet airplanes in the future are summarized. The SST using afterburning turbojets and no sound suppression is estimated to produce 129 EPNdb on the sideline. Suppression devices and operating techniques can reduce this noise markedly. The present fleet of 707's and DC-8's produces noises nearly as high as 120 decibels on approach. Combinations of sound absorbing devices and engine modifications are expected to give a noise near 100 decibels. The new advanced technology transports (ATT) are being studied with two noise goals in mind: one is 10 decibels and the other is 20 decibels below the FAR level of 106 on takeoff. New trijets using new quiet engines are expected to meet 90 EPNdb. Future advances in technology are expected to yield another 5 decibels to give noise levels of 80 EPNdb. The STOL airplane has a noise goal of 95 EPNdb along a sideline 500 feet from the runway.
Twin-jet afterbody models were investigated by using two balances to measure separately the thrust minus total drag and the afterbody drag at Mach numbers of 0.0 and 0.50 to 2.20 for a constant angle of attack of 0. Translating shroud cone plug nozzles were tested at dry and maximum afterburning power settings with a high-pressure air system used to provide jet total-pressure ratios up to 20.0. Two nozzle lateral spacings were studied by using afterbodies with several interfairing shapes. The close- and wide-spaced afterbodies had identical cross-sectional area distributions when similar interfairings were installed on each. Nozzle cant angles of -5, 0, and 5 degrees were investigated. The results show that the highest overall performance was generally obtained with the close-spaced afterbody, basic interfairings (no base), and uncanted nozzles.
A parametric study of low-boom supersonic transport airplanes with conventional configurations was made to identify the features of specific configurations that promise relatively low sonic boom overpressures (less than 47.9 N/sq m). The range of values considered was gross weight from 28,300 to 170,000 kg; cruise Mach numbers of 2 to 3.2; and wing loadings of 1436, 2870, and 4309 N/sq m. Fuselage length was varied from 49.1 to 102.4 m and fuselage diameter from 2.75 to 3.98 m. A nominal Mach 2 configuration weighing 56,700 kg and having a wing loading of 2870 N/sq m was selected; and its gross geometric, aerodynamic, and structural features were estimated. At a cruise altitude of 18,300 m, lift-drag ratio was estimated to be 7.35, while sonic boom overpressure was 41.7 N/sq m. Takeoff thrust loading using four afterburning turbojet engines at maximum dry thrust was 0.32. Payload for a 4440-km range was 16.7 percent of gross weight, giving a direct operating cost of 0.82 cent per seat statute mile.
The dispersion of the exhaust of turbojet engines into the atmosphere is estimated by using a model developed for the mixing of a round jet with a parallel flow. The analysis is appropriate for determining the spread and dilution of the jet exhaust from the engine exit until it is entrained in the aircraft trailing vortices. Chemical reactions are not expected to be important and are not included in the flow model. Calculations of the dispersion of the exhaust plumes of three aircraft turbojet engines with and without afterburning at typical flight conditions are presented. Calculated average concentrations for the exhaust plume from a single engine jet fighter are shown to be in good agreement with measurements made in the aircraft wake during flight.
Measurements of engine noise generated by an F-111A airplane positioned on a thrustmeasuring platform were made at angles of 0 deg to 160 deg from the aircraft heading. Sound power levels, power spectra, and directivity patterns are presented for jet exit velocities between 260 feet per second and 2400 feet per second. The test results indicate that the total acoustic power was proportional to the eighth power of the core jet velocity for core exhaust velocities greater than 300 meters per second (985 feet per second) and that little or no mixing of the core and fan streams occurred. The maximum sideline noise was most accurately predicted by using the average jet velocity for velocities above 300 meters per second (985 feet per second). The acoustic power spectrum was essentially the same for the single jet flow of afterburner operation and the coaxial flow of the nonafterburning condition. By varying the inlet geometry and cowl position, reductions in the sound pressure level of the blade passing frequency on the order of 15 decibels to 25 decibels were observed for inlet Mach numbers of 0.8 to 0.9.
A computer simulation of the YF-12 aircraft motions and propulsion system dynamics is presented. The propulsion system was represented in sufficient detail so that interactions between aircraft motions and the propulsion system dynamics could be investigated. Six degree-of-freedom aircraft motions together with the three-axis stability augmentation system were represented. The mixed compression inlets and their controls were represented in the started mode for a range of flow conditions up to the inlet unstart boundary. Effects of inlet moving geometry on aircraft forces and movements as well as effects of aircraft motions on the inlet behavior were simulated. The engines, which are straight subjects, were represented in the afterburning mode, with effects of changes in aircraft flight conditions included. The simulation was capable of operating in real time.
Twin-jet afterbody models were investigated by using two balances to measure separately the thrust minus total axial force and the afterbody drag at Mach numbers from 0 to 1.3. Angle of attack was varied from minus 2 deg to 8.5 deg. Translating shroud cone plug nozzles were tested at dry-power and maximum-afterburning-power settings with a high-pressure air system used to provide jet total-pressure ratios up to 9.0. Two nozzle lateral spacings were studied by using afterbodies with several interfairing shapes. The close- and wide-spaced afterbodies had identical cross-sectional area distributions when similar interfairings were installed on each. The results show that the highest overall performance was obtained with the close-spaced afterbody and basic interfairings. Increasing angle of attack decreased performance for all configurations and conditions investigated.
The effect of swirling flow on jet noise suppression was experimentally investigated in a relatively small, low-thrust, fan-jet engine. Measurements of acoustic properties of the near and far fields, jet-flow characteristics, and engine thrust were made with and without stationary swirl vanes installed in the primary exhaust nozzle. Preliminary test results indicate that substantial reductions in jet overall sound pressure levels and overall acoustic power were obtained with minimal thrust losses. Based on preliminary analysis, present results, and previous experiments with swirling hot jets, it is predicted that even greater jet noise reductions can be obtained in higher thrust engines, particularly with afterburning, by swirling jet exhaust.
A review is presented of recent experimental results, analytical procedures and test techniques employed to evaluate the effects of inlet flow distortion on the stability characteristics of representative afterburning turbofan and turbojet compression systems. Circumferential distortions of pressure and temperature, separately and in combination are considered. Resulting engine sensitivity measurements are compared with predictions based on simplified parallel compressor models and with several distortion descriptor parameters.