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Turbojet Combustor Performance with Natural Gas Fuel

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Turbojet Combustor Performance with Natural Gas Fuel NASA TECHNICAL NOTE NASA__ TN- D-5571 z bn "COPY: REmTo AFW'L (WLOL) KIRl"D AFB, N MEII: TURBOJET COMBUSTOR PERFORMANCE WITH NATURAL GAS FUEL @~: < .' . .> 1 by Nicholas R. Marchionna and Arthur M. Trout 12 I- Lewis Research Center L CZeueZand, Ohio - .f NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. FEBRUARY 1970 TECH LIBRARY KAFB, NM 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. NASA TN D-5571 I 4. Title ond Subtitle 5. Report Date TURBOJ E T COMBUS TOR PERFORMANCE February 1970 WITH NATURAL GAS FUEL 6. Performing Organizotion Code 7. Authods) 8. Performing Organization Report No. Nicholas R. Marchionna and Arthur M. Trout E-5204 IO. Work Unit No. 9. Performing Organizotion Name ond Address 720-03 Lewis Research Center 11. Contract or Grant No. National Aeronautics and Space Administration Cleveland, Ohio 44135 13. Type of Report and Period Covered 12. Sponsoring Agency Name ond Address Technical Note National Aeronautics and Space Administration Washington, C. 20546 D. 14. Sponsoring Agency Code 115. Supplementary Notes 16. Abstract Natural gas fuel for advanced turbojet combustor applications was investigated with two swirl-can modular-combustors. In tests at a combustor inlet pressure of 45 psia 2 (31 N/cm ), inlet temperatures of 540' and 1140' F (556 and 889 K) and reference vel- ocities up to 190 ft/sec (57.9 m/sec). The efficiency of both combustors was near 100 percent for inlet air temperatures of 1140' F (889 K). However, at inlet air tem- peratures of 540' F (556 K), the efficiency of one design decreased rapidly with decreas- ing fuel-air ratio. Acoustical instabilities, encountered with both combustors as initiall: designed, disappeared when the blockage at the module exit plane was sufficiently in- creased. Altitude blowout and ignition performance was poorer than that of a JP-fueled combustor of similar design. Vitiated preheat inlet air had an important adverse effect on combustion efficiency. 18. Distribution Statement Unclassified - unlimited 19. Security Clossif. (of this report) 20. Security Clossif. (of this poge) 21- No. of Pages 22. Price* Unclassified Unclassified 1 34 1 $3.00 TURBOJET COMBUSTOR PERFORMANCE WITH NATURAL GAS FUEL by Nicholas R. Marchionna and Arthur M. Trout Lewis Research Center SUMMARY Natural gas as a fuel for advanced turbojet combustor applications was investigated in tests with two swirl-can modular-combustor designs. The combustors from the dif- fuser inlet to the exhaust nozzle were 34 inches (86 cm) long. The diffuser was 15 inches (38 cm) long with a 33' included angle. Burning length in the combustors was approxi- mately 20 inches (51 cm). Tests were conducted over a range of fuel-air ratios at a pressure of 45 psia (31 N/cm 2 ), combustor inlet temperatures of 540' and 1140' F (556 and 889 K) and combustor reference velocities up to 190 feet per second (57.9 m/sec). With an inlet air temperature of 1140' F (880 K) the efficiency of both combustors was near 100 percent over a range of fuel-air ratios. However, with an inlet air tem- perature of 540' F (556 K), the efficiency of one design decreased rapidly with decreas- ing fuel-air ratio. Altitude blowout and ignition performance of both combustors was poorer than that of a JP-fueled combustor of similar design. Acoustical instabilities were encountered in both gas-fueled combustor designs. The use of absorbing liners only partially dissipated the instabilities. The instabilities did disappear when the block- age at the module exit plane was significantly increased. Vitiated preheat inlet air was found to have an important adverse effect on combustion efficiency. INTRODUCTION Recent studies have shown that liquefied natural gas fuel offers significant advantages over JP-type fuels in some turbojet engine applications (e. g., refs. 1 to 3). For exam- ple, a 31-percent payload improvement was calculated for a Mach 3 supersonic transport application (ref. 1) assuming the increased heat sink capacity of the liquefied natural gas was used to allow higher turbine inlet temperatures. Alternatively, the extra heat sink capacity could be used to maintain lower turbine metal temperatures and thereby improve turbine life and reliability. Other advantages of liquefied natural gas fuel include a I- higher heating value than JP-fuels, less tendency to smoke, lower flame radiation and greatly reduced tendency to fuel decomposition (ref. 3). The use of liquefied natural gas fuel introduces some special considerations. Al- though stored as a liquid on board the aircraft, it will enter the combustor as a gas. Ex- perience with gaseous fuels in turbojet combustors is limited. A study of the results of early investigations (refs. 4 to 8) has shown that, while higher combustion efficiencies could usually be obtained with gaseous fuels, the way in which the fuel was injected had a major influence on the efficiency and on the lean and rich stability limits. This influence was most pronounced at lower pressures. Liquefied natural gas (approx. 90 percent me- thane), having narrower flammability limits than propane or other commonly used gas- eous fuels (ref. 9) may be more critical in this regard. The tests described in refer- ence 8, comparing natural gas to propane in an annular turbojet combustor, do show a narrowing of the limits of stable operation when natural gas was used. This appeared to be due partly to the poorer burning characteristics of natural gas and partly to the change in fuel distribution caused by the difference in density between natural gas and propane. When the fuel nozzle orifice size was enlarged for the natural gas, to match the momen- tum of the propane at the same fuel flow, performance was substantially improved. It can be expected from the limited information available in the literature that some problems might be experienced in getting adequate efficiency and stability range in a natural-gas-fueled turbojet combustor. Therefore, a project was undertaken to test such a combustor over a range of operating conditions. An advanced turbojet combus- tor was designed for the use of natural gas. A modular combustor design was used which had been developed in the past for gaseous fuels (refs. 10 to 12). A similar mod- ular approach was also used recently with liquid JP-fuel in a high temperature advanced combustor design (ref. 13). The primary purpose of the present study was to deter- mine the performance problems of an advanced design turbojet combustor operating on natural gas fuel. TEST INSTALLATION A 12- by 30-inch (30.5- by 76.2-cm) test section, housing the combustor array, was installed in a closed-duct test facility (fig. 1) connected to the laboratory air supply and 2 exhaust systems. Combustion air at pressures up to 150 psia (103.5 N/cm ) was passed through a nonvitiating preheater which was capable of heating the air to 600' F (589 K). For those conditions requiring a combustor-inlet temperature of 1140' F (889 K), the air was preheated further by a vitiating preheater consisting of 10 571 single combustor cans. A set of baffles was installed downstreamof the 571 cans to ensure a uniform temperature profile at the combustor inlet. Airflow rates and combustor pressures were regulated by remotely controlled valves upstream and downstream of the test section. 2 Laboratory air supply \ I I L- Air orifice J Air control valve / iitiating preheater Test r Exhaust I control I valve J.....s.,., ..... Combustor I exit plane Atmospheric or I altitude exhaustJ Figure 1. - Test facility and auxiliary equipment. W Acoustic liner 7 Film-cooled liner \ \ r33° Included \, '1 0.50 angle diffuser \ \\ (1.27) ,-Exhaust Equally spaced \ nozzle \ flow dividers -, 15.0 (38.1) 12.0 (30.5) Figure 2. - Combustor installation in test section. (Dimensions are in inches (cm).) The test sections (fig. 2) were scaled to simulate a 90' sector of a full annulus of a turbojet engine combustor with a 57-inch (145-cm) diameter outer casing, a length of 34 inches (86 cm), and a duct height of 12 inches (30 cm). For ease of fabrication, the test sections were made rectangular in cross section. The diffuser had a 33' included angle and was 15 inches (38 cm) long. Because of the steep diffuser angle, equally spaced flow dividers were installed to provide a uniform velocity profile and to prevent flow separation at the walls. The inlet sections of the combustor modules were located at the end of the diffuser. A film-cooled liner extending from the downstream end of the combustor modules to the exhaust nozzle was used to protect the outer housing. A de- scription of the test instrumentation is presented in the appendix. TEST COMBUSTORS The combustors designed for the natural gas tests were based on the modular ap- proach introduced in reference 12. In this approach, the combustor consists of an array of small swirl-can combustor modules. Each module acts as an individual flameholder. Two module designs were used in the present tests: model I, a 2-inch (5.08-cm) di- ameter module (fig. 3); and model II, a 3.5-inch (8.9-cm) nominal diameter scalloped module (fig. 4). The original version of the model I combustor array is shown in figure 5. The two strips between two modules in adjacent rows were to assist in cross-firing. The four half-modules shown in figure 5 were nonburning and were installed to provide uniform blockage. The original version of the model 11 combustor is shown in figure 6. In both models, the swirl-cans were supported by the fuel supply manifold. Fuel in- jection holes were designed so that the fuel would be injected tangentially at sonic veloc- ity at most operating conditions, assuring even fuel distribution through each row.
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