NASA TECHNICAL NOTE
CO
Io
NAASA TN )-7135
COMPARISON OF COMBUSTIONCHARACTERISTICS OF ASTM A-l,PROPANE, AND NATURAL-GAS FUELSIN AN ANNULAR TURBOJET COMBUSTOR
by Jerrold D. Wear and Robert E. Jones
Lewis Research Center
I Cleveland, Ohio 44135
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C. • JANUARY 1973
https://ntrs.nasa.gov/search.jsp?R=19730008044 2018-05-22T19:18:08+00:00Z
1. Report No.
NASA TN D-7135
2. Government Accession No. 3. Recipient's Catalog No.
4. Title and Subtitle
COMPARISON OF COMBUSTION CHARACTERISTICS OFASTM A-l, PROPANE, AND NATURAL-GAS FUELS IN ANANNULAR TURBOJET COMBUSTOR
5. Report Date
January 19736. Performing Organization Code
7. Author(s)
Jerrold D. Wear and Robert E. Jones8. Performing Organization Report No.
E-7078
9. Performing Organization Name and Address
Lewis Research CenterNational Aeronautics and Space AdministrationCleveland, Ohio 44135
10. Work Unit No.
501-2411. Contract or Grant No.
12. Sponsoring Agency Name and Address
National Aeronautics and Space AdministrationWashington, D.C. 20546
13. Type of Report and Period Covered
Technical Note14. Sponsoring Agency Code
15. Supplementary Notes
16. Abstract
This report compares the performance of an annular turbojet combustor using natural-gas fuelwith that obtained using ASTM A-l and propane fuels. Propane gas was used to simulate opera-tion with vaporized kerosene fuels. The results obtained at severe operating conditions andaltitude relight conditions show that natural gas is inferior to both ASTM A-l and propane fuels.Combustion efficiencies were significantly lower and combustor pressures for relight werehigher with natural-gas fuel than with the other fuels. The inferior performance of natural gasis shown to be caused by the chemical stability of the methane molecule.
17. Key Words (Suggested by Author(sl)
Jet engine; Combustors; ASTM A-l; Gaseousfuel nozzles; Natural gas; Propane; Combus-tion efficiency; Altitude ignition and blowout
18. Distribution Statement
Unclassified - unlimited
19. Security dassif. (of this report)
Unclassified20. Security Classif. (of this page)
Unclassified21. No. of Pages
2222. Price*
$3. 00
* For sale by the National Technical Information Service, Springfield, Virginia 22151
COMPARISON OF COMBUSTION CHARACTERISTICS OF ASTM A-l, PROPANE,
AND NATURAL-GAS FUELS IN AN ANNULAR TURBOJET COMBUSTOR
by Jerrold D. Wear and Robert E. Jones
Lewis Research Center
SUMMARY
This report compares the performance of an annular turbojet combustor usingnatural-gas fuel with that obtained using ASTM A-l and propane fuels. Propane gas wasused to simulate operation with vaporized kerosene fuel. The combustion efficiency dataobtained with these fuels is compared at several simulated off-design engine operatingpoints. These points were chosen to illustrate the differences in performance obtainablewith the three fuels. In addition, both altitude relight and combustor blowout data arecompared for the three fuels.
These investigations show that the use of natural-gas fuel results is significantlylower values of combustion efficiency at severe operating conditions, higher values ofcombustor pressure (lower flight altitude) for altitude ignition and blowout, and astronger tendency for combustion instability than either ASTM A-l or propane fuel. Theinferior performance obtained with natural-gas fuel is explained in terms of the chemi-cal stability of the methane molecule. Physical and chemical properties of the threefuels are tabulated and compared to illustrate the relative chemical stability of eachfuel.
INTRODUCTION
This report compares the combustion performance of ASTM A-l fuel and natural-gas fuels in a combustor designed for an advanced supersonic flight engine. Propanefuel is also compared as a gaseous fuel representative of vaporized kerosene fuels. Thecomparisons are made on the basis of combustion performance at off-design and altituderelight conditions, and this performance is related to fundamental combustion propertiesof each fuel.
The use of liquefied natural gas as the fuel for engines powering a supersonic trans-port has been shown to have many potential advantages over the conventional kerosenefuels (refs. 1 to 4). The more important of these potential advantages are the increasedheat-sink capability of liquefied natural gas, higher heating value on a weight basis, lowflame radiation and low smoke levels in engine exhaust. As a result of this interest innatural-gas fuel, many combustor programs were conducted to document the perform-ance attainable with natural-gas fuel (refs. 5 to 11). These programs included combus-tors designed specifically for natural-gas fuel as well as combustors designed for usewith kerosene fuel (ASTM A-l). As expected, combustor performance with natural-gasfuel was equal to that obtained with ASTM A-l fuel at combustor conditions simulatingtakeoff and cruise operation. However, combustor performance at off-design conditionswas considerably poorer with natural-gas fuel. Combustion efficiency decreased mark-edly with decreasing pressure and was particularly sensitive to a decrease in the inlet-air temperature. Of particular importance were the very poor altitude blowout and re-light limits obtained with natural-gas fuel. For every operating condition, the measuredblowout and relight pressures were significantly higher than those obtained withASTM A-l fuel(ref. 5).
A recent investigation (ref. 12) was conducted to determine if combustor perform-ance with natural-gas fuel could be significantly improved by determining the optimummethod of fuel injection. This study was deemed necessary because many previous in-vestigations (e. g., refs. 13 to 16) had indicated that the method of gaseous fuel injectionwas of primary importance in determining combustor performance. The best injectordesign for natural-gas fuel (ref. 12) was one that injected the fuel in discrete jets at ashallow angle relative to the combustor centerline. This injector also gave very goodperformance with propane fuel.
This report extends the effort of reference 12 by comparing the combustor perform-ance with natural gas, propane, and ASTM A-l liquid fuel. Propane fuel was used tosimulate vaporized kerosene fuel injected as a gas instead of as a liquid. Three differ-ent configurations of gaseous fuel nozzles were used for the comparisons.
Nominal test conditions used for combustion efficiency determinations were as fol-lows: inlet pressure, 13. 8 and 17. 2 newtons per square centimeter (20 and 25 psia);combustor reference velocity, 32. 3 and 40. 5 meters per second (106 and 133 ft/sec);and inlet-air temperature, 422 K (300° F).
The altitude relight and blowout test conditions included two combustor referenceMach numbers, 0. 08 and 0. 10, and two inlet-air temperatures, 300 and 425 K (80° and305° F).
The U. S. customary system of units was used for primary measurements and calcu-lations. Conversion to SI units (Systems International d'Unites) is done for reportingpurposes only. In making the conversion, consideration is given to implied accuracy andmay result in rounding off the values expressed in SI units.
APPARATUS
The combustor used in these tests was an advanced annular combustor described inreferences 17 and 18. This combustor was designed for use with liquid fuel and a mod-ification to the fuel injectors was necessary for use with natural gas and propane fuels.Figure 1 is a cross-sectional sketch of the combustor test section showing inlet and out-let ducting and instrumentation planes. Pertinent dimensions are included.
Figure 2 is a cross-sectional sketch of the combustor headplate showing the, liquidfuel dual-orifice fuel nozzle. The air swirler screws onto the fuel strut and acts as aretainer for the fuel nozzle. A photograph of the fuel nozzle installed in the fuel strutis also shown. Figure 3 shows the various gaseous fuel nozzles used in this study, twoof which were used in tests reported in reference 12. To provide the increased injectionarea required for the gaseous fuels, the injection plane of the nozzle was located fartherdownstream than the injection plane of the liquid fuel dual-orifice nozzle.
Nozzle 2 (fig. 3(a)) provided angled injection of the fuel through six holes with a totalo
injection area of 1. 068 square centimeters (0. 1656 in. ). This nozzle was tested withgaseous propane (data reported herein) and with natural gas (data reported in ref. 12).
Nozzle 9A is shown in figure 3(b). The injection plane is farther downstream andthe physical size of the nozzle has been substantially increased compared with nozzle 2.Angle injection through six holes with a total injection area of 1. 254 square centimeters
o(0. 1944 in. ) is provided by nozzle 9A.
Nozzle 8 (fig. 3(c)) was tested with natural-gas fuel; these data are reported inreference 12. There are 10 angled injection holes per nozzle with a total injection area
oof 3. 576 square centimeters (0. 5542 in. ). This nozzle is similar in physical size to9A; however, nozzle 8 has a larger injection area than does nozzle 9A.
The capabilities of the facility used in this investigation are given in detail in refer-ences 17 to 19.
Fuels
Chemical and physical properties of the natural gas, propane, and ASTM A-l fuelsare presented in table I. The natural-gas composition reported is representative of thatused during the test program, which was obtained from the natural gas supplied to theLewis Research Center for general use. The gas composition varied slightly and wasdependent on the seasonal demand and gas field from which it was obtained. The propanefuel was obtained from a commercial supplier. The ASTM A-l was obtained from asource that is used by commercial airlines.
Instrumentation
Combustion air flow rates were measured by square-edge orifice plates installedaccording to ASME specifications. Liquid fuel flow was measured by turbine flow metersusing frequency-to-voltage converters for readout and recording.
Combustor inlet-air total and static pressures were measured at the plane of thediffuser inlet (station 3, fig. 1). Combustor exhaust or outlet total and static pressuresand total temperatures were measured at the turbine inlet plane (station 5, fig. 1).Combustor exhaust total pressures and temperatures were measured at 3 incrementsaround the exhaust circumference. At each point, five temperature and pressure read-ings were obtained across the radius.
Exhaust thermocouples were platinum - 13-per cent-rhodium/platinum and were ofthe high-recovery aspirating type. The indicated readings of all thermocouples weretaken as true values of the total temperatures. More detail of the instrumentation con-struction, dimensions, and readout capability are given in references 17 to 19.
PROCEDURE
Combustion Efficiency Tests
Table II presents the three operating conditions used for combustion efficiency com-parisons of the nozzles. The table includes inlet pressures, inlet temperatures, massflows, reference velocities, and values of a correlating parameter PT/V. The PT/Vparameter is calculated from inlet total pressure, inlet total temperature, and combus-tor reference velocity. The different operating conditions are designated as conditions1, 2, and 3. The severity of the combustor inlet conditions in terms of PT/V increasesfrom condition 1 to 3.
Conditions 1 and 2: change in reference velocity at the same inlet pressure.Conditions 2 and 3: variation in inlet pressure at the same reference velocity.Conditions 1 and 3: constant air flow with variation in inlet pressure and reference
velocity.The procedure followed at each condition was that after ignition the inlet conditions
of pressure, temperature, and air flow were adjusted to desired values. Data weretaken at several fuel-air ratio values with 0.008 and 0. 020 being arbitrarily selected asthe lean and rich fuel-air ratio limits, respectively.
Altitude Limit Tests
Altitude limit data were taken to determine the combustor pressures where the com-bustor flame blew out and where ignition occurred. These tests were conducted as fol-lows: after ignition at a pressure considerably higher than the possible blowout pres-sure, values of inlet-air temperature and reference Mach number were held constantwhile decreasing the inlet total pressure. Fuel-air ratio was held at about 0.010 duringthe change in inlet-air pressure. At each inlet condition fuel flow was then increased toa value intended to give an approximate 556 K (1000° F) theoretical temperature rise.The fuel-flow increase was over a time period of 6 to 8 seconds. If the monitored ex-haust temperatures showed an increase during the fuel-flow increase, the fuel-air ratiowas reduced back to about 0.010, and the series of steps was repeated at successivelylower pressure levels. This procedure was repeated until combustor blowout was en-countered during the increase in fuel flow.
Pressure values for relight were determined as follows: at the desired inlet condi-tions, the fuel-air ratio was slowly varied up and down from about 0. 005 to 0. 015 (dur-ing a maximum time period of 60 sec). If ignition occurred and combustion was stableat this fuel-air ratio, the inlet pressure was recorded as an ignition pressure.
CALCULATIONS
Combustion Efficiency
Efficiency was determined by dividing the measured temperature rise across thecombustor by the theoretical temperature rise. Exit temperatures were measured withfive-point, traversing, aspirated thermocouple probes and were mass-weighted for theefficiency calculation. The inlet temperature was the arithmetic average of readings ofeight single-point thermocouples around the inlet circumference. The theoretical tem-perature rise was computed as a function of fuel (heat of formation and hydrogen-carbon
.weight ratio), inlet-air pressure, inlet-air temperature, and fuel-air ratio.The composition of the natural gas as shown in table I indicated about 97 to 98 per-
cent hydrocarbons. The heating value and fuel-air ratios used for theoretical tempera-ture rise and other calculations and figures were based on actual hydrocarbons in thegas. The nonhydrocarbons were considered to be air.
Inlet-Air Total Pressure
The average inlet-air total pressure was obtained by mass-weighting values fromeight five-point pressure rakes around the diffuser inlet. Static pressures, used in the
mass weighting calculations, were measured around the circumference on both the innerand outer wall of the inlet annulus.
Combustor Reference Mach Number and Velocity
The combustor reference Mach number was computed from the total air flow, inlettotal pressure and temperature, and reference area (maximum cross-sectional area be-otween inner and outer shrouds, 4484 square centimeter (695 in. )).
Reference velocity for the combustor was computed from combustor reference Machnumber and sonic velocity at the particular inlet condition.
RESULTS AND DISCUSSION
Combustion Efficiency Tests
A summary of combustion efficiency test data with ASTM A-l and propane is tabu-lated in table m. Figure 4 compares the combustion efficiency measurements obtainedwith the three fuels. The combustion efficiency with propane and natural gas fuels isslightly better than that obtained with ASTM A-l fuel at the milder operating conditionsof figure 4(a). Also shown is the efficiency of natural-gas fuel with nozzle 8. Unstablecombustion occurred at a fuel-air ratio in excess of 0. 019 with this nozzle, and combus-tion efficiency was considerably below that obtained with nozzle 2 with natural-gas fuel.The comparison in combustion efficiency between propane in nozzle 9A and natural gasin nozzle 8 reflects differences in the combustion properties of the two fuels. Thesenozzles were designed so that the gaseous fuels were injected at the same velocity forsimilar weight flow rates. Therefore, injection velocity effects cannot account for thelarge differences in efficiency. Injection velocity effects may contribute to the combus-tion efficiency differences between propane and natural gas obtained with nozzle 2. Fig-ures 4(b) and (c) show the effects of increasingly severe operating conditions on combus-tion efficiency. In each case the efficiency with propane fuel exceeds that of ASTM A-lfuel. Natural-gas fuel combustion efficiency decreases markedly with unstable combus-tion frequently occurring.
The data are replotted in figure 5 with combustion efficiency shown against the com-bustion parameter PT/V for three values of heat content per unit weight of air. Foreach fuel the fuel-air ratio, which provides the desired heat content per unit weight ofair, was calculated (see table IV). The corresponding values of combustion efficiencywere then obtained from figure 4. The combustion efficiency of natural-gas fuel rapidlydecreases with increasing test condition severity. Only at the mildest operating condi-
6
tion does natural-gas efficiency equal or exceed that of ASTM A-l fuel. As before, theefficiency of propane fuel is better than that of ASTM A-l fuel.
Effect of fuel injector design. - Reference 12 covers work on a wide variety of fuelinjectors for use with natural-gas fuel. The best injector of that work (injector 2) wasused in this program. Comparing the combustion efficiency obtained with this injectorindicates that natural-gas fuel is inferior to propane as a fuel. There has been a largebody of work done on gaseous fuel injectors, and the results of those tests confirm theconclusion that no method of injecting natural-gas fuel yet tested will give combustionefficiency as good as that obtained with propane fuel in the same injector. Many injec-tors have given good efficiency but none have given combustion efficiency equal to thatobtained with propane fuel or JP-type kerosene fuel at these severe operating conditions.
Effect of common injection velocity. - The data shown in figure 4 also compare thecombustion efficiency for natural gas and propane fuels at the same injection velocity.At the same fuel weight flow rate, fuel nozzle 8 with natural gas and fuel nozzle 9A withpropane inject the fuel at the same velocity. Propane fuel again gives clearly superiorcombustion efficiency at every fuel-air ratio. The differences seem not to be relevantto the nozzle design, but rather to the fuel itself.
Effect of fuel properties. - Table V is a compilation of various physical and com-bustion properties for methane (natural gas), propane, and ASTM A-l fuel. A compari-son of the properties of propane and ASTM A-l indicate that at least as far as the moreimportant combustion properties are concerned, propane is a fair representation ofvaporized ASTM A-l fuel. An examination of the properties of methane (natural gas)indicate that it is a stable hydrocarbon with narrow combustible limits and high diffu-sivity in air. The following properties indicate why performance with natural gas isconsistently inferior to the other fuels at severe operating conditions. The narrow com-bustible limits indicate that combustion can occur only over a limited range of fuel-airratios in the primary zone. This in turn requires critical design to optimize combustionintensity over the wide ranges of fuel and air flows typical of turbine engine combustors.The low molecular weight and hence high diffusivity of methane mean that the fuel quicklydisperses in the turbulent regions of the primary zone, and fuel-air ratios can quicklyfall below the combustible limit. This observation is supported by results of unreportedinvestigations of emission measurements, at low efficiency off-design operation, con-ducted at NASA. Carbon monoxide did not appear in the exhaust gas samples, whichimplies that virtually none of the inefficiency was caused by partial combustion or oxida-tion of the fuel. Other properties listed such as bond strength, and spontaneous ignitiontemperature, point to the basic chemical stability of the methane molecule.
These factors explain why poor combustion efficiency was obtained with natural-gasfuel at the severe operating conditions. As previously mentioned, the combustor usedin these tests was designed for use with ASTM A-l liquid fuel. In order to optimizecombustor performance with natural gas, combustor. modifications or redesign will be
required. Design changes may include an increase in the number of fuel injection points,an increase in the capability to vary the primary air flow, and changes in combustorcombustion volume.
Altitude Limit Tests
A summary of altitude limit test data with ASTM A-l and propane is listed in ta-ble VI. The altitude limit test results with the three fuels are shown in figures 6 and 7.These results show the effects of fuel types, nozzle design, reference Mach number,and inlet-air temperature on the pressure where satisfactory ignition is obtained andwhere combustion blowout occurs. The ignition data are shown in figure 6. The ignitionpressures of propane and ASTM A-l fuel are markedly superior to (lower than) that ofnatural-gas fuel. Increasing the inlet-air temperature does improve the results withnatural gas especially at the lower values of reference Mach number. The combustorpressures at blowout for the three fuels are shown in figure 7. At the lower inlet-airtemperature, the blowout data with natural-gas fuel is again inferior to that of propaneand ASTM A-l. At the higher inlet-air temperature of 425 K, the differences are rela-tively minor and decrease further as the reference Mach number decreases.
The poor performance of natural-gas fuel relative to that of propane and ASTM A-lat altitude relight conditions is explainable in termse of the properties of natural gasmentioned previously. These are the narrow combustion limits, fuel stability, and highdiffusivity. The high spontaneous ignition temperature of methane is a measure of thedifficulty of igniting the fuel. The narrow combustible limits require that a near stoi-chiometric mixture of fuel exist in the area of the ignitor for a time sufficient to havecombustion initiated. This requires a careful control of fuel-air mixture near the igni-tor. Such careful control is not required with fuels having wider stability limits andlower ignition temperature.
Combustion Instability
Virtually every combustor tested using natural gas or methane fuel has encounteredconsiderable combustion instability. References 6 and 12 describe these combustors andthe difficulties encountered with combustion instability. Conversely, combustors testedusing ASTM A-l and propane have been almost entirely free of any form of combustioninstability. This characteristic of natural-gas fuel is also explainable in terms of itsnarrow combustible limits. Natural-gas combustion will not be initiated until the fuel-air ratio is within the combustible range. Once there, the gaseous fuel mixture burnsrapidly. A rapid increase in combustor temperature and bulk gas velocity then occurs
8
virtually within a single axial plane of the combustor. This situation is ideal for theonset of combustion instability. Conversely, the wider limits of combustion of propaneand ASTM A-l fuels mean that combustion can be initiated while the fuel-air mixture isvery rich. This spreads the combustion axially within the combustor, and combustioninstability has rarely been encountered.
CONCLUDING REMARKS
Tests were conducted to compare the performance of an annular turbojet combustorusing natural-gas fuel with that obtained using ASTM A-l and propane fuels. The com-bustor was designed for use with kerosene fuels.
Physical properties that make the use of natural-gas fuel attractive as a heat sinkfor future high-speed aircraft also make natural gas a poor choice as the fuel. The highthermal stability of the methane molecule so necessary when used as a heat sink makethe combustion performance with this fuel poor at severe operating conditions. Normalground starting, takeoff, and cruise conditions are relatively mild operating conditions,and performance with natural-gas fuel is comparable with kerosene fuel. However, atoff-design and severe operating conditions the performance with natural-gas fuel will beconsiderably poorer than that with kerosene-type fuels. This is particularly true of thealtitude blowout and ignition limits. The tendency for combustion instability is also con-siderably greater with natural-gas fuel than with kerosene fuels.
The design of a combustor for exclusive use of natural-gas fuel must be concernedprimarily with maintaining good combustion efficiency and stability at severe operatingconditions. Attaining altitude blowout and relight limits comparable with those of kero-sene fueled combustors will require a considerable effort.
Lewis Research Center,National Aeronautics and Space Administration,
Cleveland, Ohio, October 11, 1972,501-24.
REFERENCES
1. Weber, Richard J.; Dugan, James F., Jr.; and Luidens, Roger W.: Methane-FueledPropulsion Systems. Paper 66-685, AIAA, June 1966.
2. Whitlow, JohnB., Jr.; Eisenberg, Joseph D.; and Shovlin, Michael D.: Potential ofLiquid-Methane Fuel for Mach 3 Commercial Supersonic Transports. NASA TND-3471, 1966.
3. Joslin, C. L.: The Potential of Methane as a Fuel for Advanced Aircraft. Aviationand Space: Progress and Prospects. ASME, 1968, pp. 351-355.
4. Esgar, Jack B.: Cryogenic Fuels for Aircraft. Aircraft Propulsion. NASA SP-259,1970, pp. 397-420.
5. Schultz, Donald F.; Perkins, Porter J.; and Wear, Jerrold D.: Comparison ofASTM-A1 and Natural Gas Fuels in an Annular Turbojet Combustor. NASA TMX-52700, 1969.
6. Marchionna, Nicholas R.; and Trout, Arthur M.: Turbojet Combustor Performancewith Natural Gas Fuel. NASA TN D-5571, 1970.
7. Fear, James S.; and Tacina, Robert R.: Performance of a Turbojet CombustorUsing Natural Gas Fuel Heated to 1200° F (922 K). NASA TN D-5672, 1970.
8. Marchionna, Nicholas R.: Stability Limits and Efficiency of Swirl-Can CombustorModules Burning Natural Gas Fuel. NASA TN D-5733, 1970.
9. Marchionna, Nicholas R.; and Trout, Arthur M.: Experimental Performance of aModular Turbojet Combustor Burning Natural Gas Fuel. NASA TN D-7020, 1970.
10. Trout, Arthur M.; and Marchionna, Nicholas R.: Effect of Inlet Air Vitiation on thePerformance of a Modular Combustor Burning Natural Gas Fuel. NASA TMX-52711, 1969.
11. Humenik, Francis M.: Conversion of an Experimental Turbojet Combustor fromASTM A-l Fuel to Natural Gas Fuel. NASA TM X-2241, 1971.
12. Wear, Jerrold D.; and Schultz, Donald F.: The Effects of Fuel Nozzle Design onthe Performance of an Experimental Annular Combustor Using Natural Gas Fuel.NASA TN D-7072, 1972.
13. McCafferty, Richard J.: Vapor-Fuel-Distribution Effects on Combustion Perform-ance of a Single Tubular Combustor. NACA RM E50J03, 1950.
14. Smith, Arthur L.; and Wear, Jerrold D.: Performance of Pure Fuels in a SingleJ33 Combustor. in - Five Hydrocarbon Gaseous Fuels and One Oxygenated-Hydrocarbon Gaseous Fuel. NACA RM E55KD4a, 1956.
15. Norgren, Carl T.; and Childs, J. Howard: Effect of Liner Air-Entry Holes, FuelState, and Combustor Size on Performance of an Annular Turbojet Combustor atLow Pressures and High Air-Flow Rates. NACA RM E52J09, 1953.
16. Norgren, Carl T.; and Childs, J. Howard: Effect of Fuel Injectors and Liner Designon Performance of an Annular Turbojet Combustor with Vapor Fuel. NACA RME53B04, 1953.
10
17. Rusnak, J. P.; and Shadow en, J. H.: Development of an Advanced Annular Com -bustor. Rep. PWA-FR-2832, Pratt & Whitney Aircraft (NASA CR-72453), May 30,1969.
18. Wear, Jerrold D.; Perkins, Porter J.; and Schultz, Donald F.: Tests of a Full-Scale Annular Ram-Induction Combustor for a Mach 3 Cruise Turbojet Engine.NASA TN D-6041, 1970.
19. Adam, Paul W.; and Norris, James W.: Advanced Jet Engine Combustor TestFacility. NASA TN D-6030, 1970.
20. Barnett, Henry C.; and Hibbard, Robert R., eds.: Basic Considerations in theCombustion of Hydrocarbon Fuels with Air. NACA Rep. 1300, 1959.
21. Barnett, Henry C.; and Hibbard, R. R.: Fuel Characteristics Pertinent to the De-sign of Aircraft Fuel Systems. NACA RM E53A21, 1953.
22. Simon, Dorothy Martin: Flame Propagation. Ill - Theoretical Considerations ofthe Burning Velocities of Hydrocarbons. J. Am. Chem. Soc., vol. 73, no. 1,Jan. 1951, pp. 422-425.
23. Sherwood, Thomas K.: Absorption and Extraction. McGraw-Hill Book Co., Inc.,1937, pp. 18-19.
24. Weast, Robert C., ed.: Handbook of Chemistry and Physics. Forty-fifth ed.,Chem. Rubber Pub. Co., 1964-1965, F-94.
25. Calcote, H. F.; Gregory, C. A., Jr.; Barnett, C. M.; and Gilmer, Ruth B.:Spark Ignition - Effects of Molecular Structure. Ind. Eng. Chem., vol. 44,no. 11, Nov. 1952, pp. 2656-2660.
26. Maxwell, J. B.: Data Book on Hydrocarbons. D. Van Nostrand Co., Inc., 1950,pp. 10-21.
27. Hibbard, Robert R.: Evaluation of Liquified Hydrocarbon Gases as Turbojet Fuels.NACA RM E56121, 1956.
11
TABLE I. - CHEMICAL AND PHYSICAL PROPERTIES OF FUELS
(a) Natural gas and gaseous propane
Density, a kg/m3 (lb/ft3)Net heat of combustion (calculated)
J/kg (Btu/lb)Normalized chromatographic
analysis, vol. %:MethaneEthanePropaneC,, Cg, and Cg hydrocarbonsNitrogenCarbon dioxideOxygen
Natural gas
0.7320(0.0457)49 770X103 (21 397)
93.503.530.530.321.051.07
trace
Gaseous propane
1.8646 (0. 1164)46 315X103 (19 925)
0. 150. 17
99.610.030.03
0.01
(b) ASTM A-l
Gravity, °API (D287)ASTM distillation (D86), K (°F):
Initial boiling point5 Percent evaporated10 Percent evaporated30 Percent evaporated50 Percent evaporated70 Percent evaporated90 Percent evaporated95 Percent evaporatedFinal boiling point
Residue, percentLoss, percentFlash point (D56), K (°F)Pour point (D97), K (°F)Viscosity at 239 K (-30° F)(D445), ra2/sec (cS)Aromatics (D1319), vol. %Net heat of combustion (D1405), JAg (Btu/lb)
43.1
433 (320)444 (340)455 (360)472 (390)483 (410)495 (431)519 (474)533 (500)547 (525)
1.10.9
324 (124)233 (-58)
9.2X10"6 (9.2)15.51
43 270X103 (18615)
aAt 289 K (60° F) and 10.159 N/cm2 (30.00 in. Hg at 32° F).
TABLE II. - COMBUSTOR NOMINAL OPERATING CONDITIONS
(Temperature, 422 K (300° F).]
Operatingcondition
123
Pressure
N/cm
17.217.213.8
psia
25.025.020.0
Air flow rate
kg/sec
20.625.920.7
Ib/sec
45.557.045.6
Reference ve-locity
m/sec
32.340.540.5
ft/sec
106133133
PT— — ParameterV
N-K-sec
m3
22. 53X105
17.9514.36
lb-°R-sec
ft3
25.81X103
20.5716.46
12
TABLE m. - COMBUSTOR EFFICIENCY DATA
(a) Fuel, ASTM A-l
Combustor inlet-air conditions
Pressure,tota
N/cm2 psia
17.2n.o17.117.117.217.217.1
17.1n n
. £
17.1
17.1
17.0
13.7
13.6
13.6
13.7
13.7
13.6
13.8
13.7
13.71^ 7io. /
13.7
25.024.724.824.824.924.924.8
24.8nA Q£*t, V
24.8
24.824.7
19.8
19.7
19.7
19.8
19.9
19.7
20.0
19.9
19.81Q ftiy. o19.9
Temper-ature,total
K
424
423
421
423
421
422
421
432491t6Q
421423
423
421
429
42S
424
425
424
419
420419491*t6l
420
°F
304
301
299
301
299
300
298
318
301299301
301
299
312
306
303
305
303
295
296
295OQQ690
296
Flow
kg/sec B/sec
Referencevelocity
ra/sec ft/sec
Parameter PT/V
N-K-sec
m3
Ib-°R-sec
ft3
Fueltemper-
ature
K °Fr
Manlfold-combustor
fuel pressuredifferential
M//*mIN/ Cm psid
Calculatedfuel injec-
tion velocity
m/sec ft/sec
Fuel-airratio
Combustoraverage ex-haust tem-perature,
total
K °F
Combus-tor tem-perature
rise
V °F
Combus-tion effi-ciency,
percent
Dual orifice fuel nozzle
19.920.420.420.420.420.520.4
23.7nc ftZD. U
25.125.125.1
20.021.120.320.220.220.120.520.520.59n A£\J. 4
20.4
43.944.944.945.045.045.144.9
52.2fiC •>33. £i
55.455.455.3
44.046.544.844.544.544.445.345.245.145 045.0
31.432.032.032.032.032.032.0
37.839 039.339.339.3
39.042.140.239.639.639.939.639.639.639 939.6
103
105
105
105
105
105
105
124128129
129
129
128
138
132
130
130
131
130
130
130
131
130
23.35X105
22.4422.5322.54
•22.7022.6422.49
19.5118. 5618.3518.4618.32
14.7813.8514.3114.5814.6814.4514.5814.5814.4114. 4014.54
26.76X103
25.7125.8225.8326.0125.9425.77
22.3621. 2721.0321.1520.99
16.9415.8716.4016.7116.8216.5616.7116.7116.5116. 5016.66
303
295
295
296
296
298
297
317297299
298
298
298
293
295
294
294
296
291
294
294
295
295
85
72
71
74
73
76
75
1117578
76
76
77
67
71
70
70
73
6570
69
72
71
79.381.684.987.991.294.294.9
81.088. 192.695.599.4
79.379.581.585.288.491.681.585.288.391 594.4
115.0118.3123.2127.6132.3136.6137.7
117.5127 8134.3138.6144.1
115.1115.3118.1123.6128.2132.8118.2123.6128.1132. 7137.0
— -
— -
...
...
...
...
...
...
...
...
------...
—..._..
—
—...
—
0.0084.0101.0133.0162.0193.0224.0234
.00860134
.0165
.0197
.0229
.0084
.0079
.0102
.0133
.0165
.0195
.0101
.0132
.01630194
.0224
686753
870
981
109612061242
702867971
10871205
670
655
734
847
954
1055725839
935
10421161
776
895
11061306151317121776
8041101128814971709
747
720
862
106512581439846
1051122414161630
263
331
448
558
674
784
821
270444
549
664
783
249
227
309
423
530
631
306419
516
621
741
473
595
807
1005121414121478
486800989
11961409
448
409
557
762
954
1136551
755
929
11181333
79.183.388.291.494.5 '96.296.8
79.686. 688.691.694.4
75.072.477.683.385.487.777.482.784.386. 991.0
(b) Fuel, gaseous propane
Fuel nozz e 2
17.0
17.4
17.2
17.0
16.9
13.6
13.6
13.5
24.625.224.9
24.724.5
19.719.719.6
425
426
419
426
433
420421
420
305
308
295
308
319
297
298
297
20.920.922.3
24.925.1
20.520.520.5
46.046.049.1
55.055.3
45.345.345.3
33.232.634.7
39.640.5
40.240.240.5
109
107
114
130
133
132
132
133
21.74X105
22.7320.76
18.3418.05
14.2014.2414.04
24.91X103
26.0523.79
21.0220.68
16.2716.3216.09
322
329
309
303
294
303
306
303
120132
96
8569
85
91
86
25.636.314.2
16.027.2
24.116.310.0
37.252.620.5
23.139.4
34.923.614.5
33.243.023.5
26.833.2
36.329. .022.3
109
141
77
88
109
119
95
73
0.0117.0152.0079
.0080
.0102
.0107
.0084
.0063
877
1016705
713
818
809
719
635
11191370809
8231013
997
834
683
452
589
286
286386
389
298
214
814
1061514
514
694
700
536
386
93.795.884.8
83.790.3
87.083.878.5
Fuel nozzle 9A
17.217.217.217.1
17.217.217.217.2
13.713.713.713.8
25.024.924.924.8
25.024.924.924.9
19.819.919.920.0
426
428432
434
425
431434
433
426
427
433
435
308
310
318322
305
316322
320
307
309
320
323
21.121.121.321.5
26.426.326.226.3
21.121.121.521.5
46.546.646.947.3
58.158.857.858.0
46.546.547.347.3
33.233.534.134.7
41.141.842.142.1
41.841.542.742.7
109
110
112
114
135
137
138
138
137
136
140
140
22.14X105
21.8521.7321.38
17.8317.7617.7817.66
13.9614.1013.9414.05
25.37X103
25.0424.9024.50
20.4320.3520.3720.24
16.0016.1615.9116.10
291
295
308302
292
301
319
305
291
284
324
326
6572
9584
66
82
114
90
64
52
124
127
42.633.720.027.2
45.327.236.156.2
37.047.223.532.5
61.848.929.039.5
65.639.452.481.6
53.768.534.147.1
36.932.323.228.3
38.429.037.544.5
39.646.031.737.8
121
106
76
93
126
95
123
146
130
151
104
124
0.0180.0149.0098.0122
.0144
.0099
.0125
.0162
.0145
.0182
.0097
.0121
11411016806
906
964
783
890
1065
938
1100762
859
15951370991
1172
1275950
11431458
12291521912
1087
715
589
374
472
539
352
456
632
512
673
329
425
12871060673
850
970
634
821
1138
922
1212592
765
100.798.491.794.3
92.985.189.197.6
87.794.481.285.2
13
TABLE IV. - HEAT CONTENT PER UNIT
WEIGHT OF AIR FOR THREE FUELS
AT VARIOUS FUEL-AIR RATIOS
Fuel
ASTM A-l
Propane
Natural gas
Fuel-airratio
0. 0090.0140.0200
.0084
.0131
.0187
.0078
.0122
.0174
Heat content
J/gair
389606865
389606865
389606865
Btu/lbair
168261372
168261372
168261372
TABLE V. - FUNDAMENTAL PHYSICAL AND COMBUSTION PROPERTIES OF
METHANE, PROPANE, AND ASTM A-l FUELS
Flammability limits, percent stoichiometric:LeanRich
Maximum burning velocity, cm/secSpontaneous ignition temperature, K (°F)
nDiffusion coefficient, cm /secChemical bond strength, kJ/mole (kcal/mole)Minimum ignition energy, JMolecular weightReaction rate at gas temperature of 750 K
(890° F), mole/sec
Methane
a46a!64
C33.8a905(1170)
eO. 181f435 (104)
h4. 70X10"4
16.04J7X10"9
Propane
a51a283
C39.0a778 (940)
eo. 100g410 (98)
h3. 05X10"4
44.10J4X10"5
ASTM A-l
b51b385
a'd516 (468)e0.050
1164
RefcRefdJeteReffRef.gRefhRef.
|Ref.jRef.
. 20,
. 21,
. 22.fuel,. 23.
24,. 24,
25.26.27.
appendix table XXXII.eqs., p. 29.
grade, JP-5.
H-CH3.H-n-C3H7.
14
TABLE VI. - COMBUSTOR ALTITUDE LIMIT DATA
Combustor inlet-air conditions
Pressure,tfifn 1IUIH1
, 9
N/cm'' psia
Temper-ature,
total
K °F
Flow
kg/sec Ib/sec
ReferenceIw3.cn nu in —
ber
Rating criteria
Ignition;stiiDic combustionat ignition fuel-
air ratio
Combustion-blowout3S IU6l~3.ll* 1*311 0 1J1~
creased above 0.01;stable combustionat fuel-air ratio of
0.010 or less
Fuel, gaseous propane; fuel nozzle 9A
5.54.84.13.43.1
9.78.36.96.2
5.5
3.42.8
2.8
5.53.4
2.8
8.07.06.05.04.5
14.012.010.09.08.0
5.04.0
4.0
8.05.04.0
298298300296297
301298300298
298
416416
416
414415
416
77
76807375
82768077
76
290290290
285288290
7.86.85.94.93.9
17. 114.712.211.09.8
4.23.4
3. 4
8.45.3
4.3
17.215.012.910.88.6
37.732.326.924.221.5
9.37.57.5
18.511.69.4
0.078.078.079.078.069
.099
.098
.098
.098
.098
.080• .081
.081
. 100
. 100
. 102
YesYesNo_..
YesYesYesNoNo
YesNo
---
YesYesNo
NoNo...
NoYes
NoNoNoNoNoa
No
—Yes
NoNoYes
Fuel, gaseous propane; fuel nozzle 2
9.79.05.54. 13.42.8
14.013.08.06.05.04.0
422422422422422422
300300300300300300
14.513.68.46.4
5.34.2
32.030.018.614.011.69.3
0. 100
.101
. 101
. 102
. 101
.101
—YesYesNoNo
---
No...--.
.NoYes
Fuel, ASTM A-l dual orifice fuel nozzle
9.77.66.25.54.84.13.4
3.3
9.76.95.54.13.42.82.1
14.011.09.0
8.07.06.05.0
4.8
14.010.08.06.05.04.03.0
305303303303303305304
304
425425424415416419420
908585
86859087
87
305305
303288290295296
17. 113.411.09.88.57.36. 15.9
14.710.58.46.35.34.23.2
37.829.624.221.518.816.013.413.0
32.523.218.613.811.79.37.0
0. 100.099.099
.099
.099
.098
.099
.100
.102
.101
.102
.099
.101
.101
.101
YesYesNoNo......
Yes..-...
YesNoNo
---
NoNoNo
—...
NoNoYes
NoNoNoNoNoNoYes
^Facility limit.
15
•o i<a 3
16
C-71-590
Original dualorifice liquidfuel nozzle-\
_6.1 cm(2.4 in.
Air swirlercenter hole;diam, 1.016cm(0.400 in.)
10.2 cm(4.0 in.)
^-Headplate
Liquid fuel
LAir swirler
rConbustor. 1 / liner
Figure 2. - Liquid fuel dual-orifice nozzle fuel strut installed in combustor head plate.
17
Fuel nozzle n
C-69-3937
Holediam, 0.476cm 10.188 in.);six holes on 1.27-cm-(0.50-in.-)diam circle-.
(a) Fuel nozzle 2; tested with gaseous propane and natural gas.
Fuel nozzle-\
T3.% cm
Q\1.56in.)
C-70-266
0.516cm
(0.203 in.); six holeson3.02-cm-)(1.19-in.-)diam circle
(b) Fuel nozzle 9A; tested with gaseous propane.
Figure 3. - Gaseous fuel nozzle installed in fuel strut.
18
C-70-263
Fuel nozzle-\
LHolediam, 0.675cm(0.266 in. I; Wholeson 3.02-cm-IL19-in.-l dian circle
(cl Fuel nozzle 85 tested with natural gas.
Figure3. - Concluded.
19
110,
100
90
70
60
50
Fuel
O ASTM A-lD Propane, gaseous
Natural gas (ref. 12)
-V Unstable combustion
Fuel nozzle
Dualorifice
I I I I I I(a) Nominal operating condition 1: Pressure, 17.2 newtons per square centimeter
(25psia); temperature, 422K(300°F); reference velocity, 32.3 meters per second(106 ft/sec).
lOOr—
90
80
.1 70
Dual—- V' orifice
60 —
I I I I I I I I50(b) Nominal operating condition 2: Pressure, 17.2 newtons per square centimeter
(25psia); temperature, 422K(300°F); reference velocity, 40.5 meters per second(133 ft/sec).
100
90
SO
70 —
60
50
40.006 .008 .010 .012 .014 .016 .018 .020 .022 .024
Fuel-air ratio
(c) Nominal operating condition 3: Pressure, 13.8 newtons per square centimeter(20 psia); temperature, 422 K (300° F); reference velocity, 40. 5 meters per second(133 ft/sec).
Figure 4. - Variation of combustion efficiency with fuel-air ratio for various fuel noz-zles and fuels.
KB,—
Fuel• ASTM A-l. Propane, gaseous- Natural gas (ref. 12)
Fuel nozzle-9A
Dualorifice
70-
(a) Heat content, 389 joules per gram of air (168 Btu/lbajr); ASTM A-lfuel-air ratio, (LOOM.
&
II3
I
-.-"11--2-.--"'" -2
to) Heat content, 606 joules per ym of air (261 Btu/lt>,ir>; ASTM A-lfuel-air ratio. 0.0140.
imp
100
orifice
I I I I I IM 16 18 20 22
PT/V parameter, N-K-sec/m3
M 18 22 26PT/V parameter, lb-°R-sec/ft3
(c) Heat content, 865 pules per gram of air (372 Btu/lt>air); ASTM A-lfuel-air ratb, 0.0200.
Figure 5. - Variation of combustion efficiency with combustion correla-ting parameter for various fuel nozzles and fuels.
20
Fuel
ASTM A-l
Fuel
ASTM A-l
16
12
8
4
0
1C
— 8
,
E< 4zto
— £ 0
— — -i_> — riupdiic, ijctteuua .Natural gas (ref. 12) 8
Fuel nozzle .o
X'
^ ^2 4
^, ̂ " Dual— ^t_ ""^-toA orifice o
4
n
U~— rrupdiie, ydieuuiNatural gas (ref. 12)
Fuel nozzle
^ - ^ "9 A
— — O — Dual orifice
1 1 1 1T_r- L_r~v« ^
J (a) Nominal inlet air temperature, 425 K (305° F).CD
1 1 1 'I^ 20£
S (a) Nominal inlet air temperature, 425 K (305° F). | |
f 20(U
_C
e .24
20
16
12
g
•SI
rf 168
— eE'E 12
— S
— so
4
r— '£ 24
/ 5 a- O>
J^^ -s 16
/* =•/ "E
,, o 19/ / 0 "
'/ 8' 04r~^
„„ ** ̂ J^— — • uuai ^,-'.'"" orifice
r ' I I „
£~ ^ 16
— i"E
o 12— to
X3
O
"~ 8
—
4^»
n
—
— /S"
_ ^
x-2Facility limit ^ ^ ^Facility limitNo blowout 7 ^ ^ /' No blowout
f ^^ JTYir*
- ^ „-*-'''.Q--'" — O — Dual orifice
1 1 1 1
/^^ — *!.„.»„, »<«»«» nA^^K*«..mh«r -07 .08 .09 .10 . 11
(b) Nominal inlet air temperature, 300 K (80° F).
Figure 6. - Minimum combustor pressures for satisfactoryignition with various fuels.
Combustor reference Mach number
(b) Nominal inlet air temperature, 300 K (80° F).
Figure 7. - Combustor blowout pressure for various fuels.
NASA-barley, 1973 24 E- 7078 21
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