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INCA RM E55G22 MTIONAL A D T T I S O R Y COMMITTEE F O R Al3RONA.UTIC S RESEARCH moRA.NDuM PRELIMJXARY EVALUATION O F FLIGHT -WIGHT X R J47 - W - 5 RAM- JET E2JGIN.E A T A MACH NUMBE3 O F 2.75 By Henry J. Welna and Dwight H. Reilly S U M M A R Y A f r e e - j e t investigation of the p e r f o m n c e , burner-shell cooling, , and ignition characteristics of a flight-weight 48-inch-diameter XRJ47- W - 5 ram-jet engine was conducted i n a Lewis a l t i t u d e t e s t chamber a t an i n l e t Mach number of 2.75 and an angle of attack of 3'. Data were ob- tained over a range of a l t i t u d e s from 58,000 t o 73,000 f e e t (engine- i n l e t a i r flows of 120 t o 60 lb/sec, respectively), a range of i n l e t temperatures from 860' t o 990' R, and a range of f u e l - a i r r a t i o s from lean blow-out t o about 0.050. The range of combustor-inlet t o t a l pres- sures f o r these conditions was from 2150 t o 960 pounds per square foot absolute.
A combustion efficiency of 0.85 was obtained a t an air flow of 60 pounds per second, i n l e t tenperature of 990' R, and the approximate de- The combustor total-pressure r a t i o at sign f u e l - a i r r a t i o of 0.043.
t h i s condition was 0.80. This combustion efficiency and combustor t o t a l - pressure r a t i o closely compare with the performance of the heavy-duty version of t h i s engine previously investigated. A t an i n l e t temperature of 990' R, flight-engine lean blow-out occurred between a f u e l - a i r r a t i o of 0.027 t o 0.028 f o r the range of a l t i t u d e investigated.
Three methods were investigated f o r mounting the cooling l i n e r s ;
t h e most promising one was used in the engine f o r approximately 3-2 hours
of burning operation without significant damage. The only consistently r e l i a b l e method of igniting t h e engine was spontaneous ignition of 1 t o 6 pounds of alurainumtrimethyl injected i n t o the engine combustor. This method was r e l i a b l e only when t h e injection r a t e exceeded 2000 pounds and when the specific gravity of the aluminum trimethyl was per hour, 0.85, INTRODUCTION The 48-inch-diameter XRJ47 -W-5 ram- j e t engine i s being developed f o r use i n the Navaho I1 missile as p a r t of the MX-770 long-range surface-to-surface missile program sponsored by t h e U . S , A i r Force.
Full-scale direct-connect and f r e e - j e t investigations of the engine a r e being conducted i n a Lewis a l t i t u d e t e s t chamber a t simulated a l t i t u d e conditions t h a t w i l l be encountered i n the cruise portion of t h e missile f l i g h t plan. The direct-connect and f r e e - j e t investigations of the heavy-duty t e s t engine have been completed and the r e s u l t s a r e presented i n references 1 and 2.
The purpose of the investigation reported herein w a s t o evaluate t h e flight-weight engine version of the XRJ47-W-5 engine. This evalua- t i o n included the determination of the performance, burner-shell cool- ing characteristics, l i n e r durability, and s t a r t i n g characteristics.
The i n t e r n a l geometry of t h e engine used i n t h i s investigation conformed t o t h e f i n a l combustor configuration arrived a t during t h e development of the combustor in the heavy-duty engine e a r l i e r in the program ( r e f . 2).
Performance data were obtained over a range of a i r flows from 120 t o 60 pounds per second, nominal i n l e t temperatures from 860' t o 990' R and f u e l - a i r r a t i o s from about 0.050 t o lean blow-out. These values of air flow, i n l e t temperature, and f u e l - a i r r a t i o correspond t o a range of combustor-inlet t o t a l pressures of 2150 t o 960 pounds per square foot absolute.
Three cooling-liner configurations were investigated. Liner length and average cooling-passage height were approximately the same f o r a l l configurations. Principal differences included the type of l i n e r corru- gations and the method of securing t h e l i n e r in the combustor. Three different ign5tion sources were evaluated. One ignition source con- s i s t e d of f l a r e s attached t o t h e flameholder, The other two sources were injection of small quantities of boron t r i e t h y l and aluminum tri- methyl i n t o the combustor i n t h e region of the flameholder.
Engine and I n s t a l l a t ion The XRJ47-W-5 ram-jet f l i g h t engine w a s i n s t a l l e d in the 14-foot- The f a c i l - diameter t e s t section of an a l t i t u d e t e s t chamber ( f i g . 1).
i t y a i r system supplied dried and heated a i r which entered the super- sonic nozzle and was accelerated t o a Mach number of 2.75 a t the entrance The operation and performance d e t a i l s of the f a c i l - of the i n l e t duct, i t y a r e presented in reference 3.
The i n s t a l l a t i o n consisted of a heavy-duty supersonic i n l e t d i f - fuser and the flight-weight conical engine diffuser, combustor, and ex- haust nozzle.
A diagram of t h e f l i g h t engine without the supersonic in- l e t diffuser i s shown in figure 2. The supersonic i n l e t diffuser was designed f o r a Mach number of 2-75; performance of t h i s diffuser is pre- sented i n reference 4. Attached t o the supersonic diffuser outlet was t h e 30' conical engine diffuser, which was 29.9 inches long and termin- ated a t t h e entrance t o the 48-inch-diameter combustor. The combustor was 60.4 inches long and was made of 0.060-inch Inconel. Affixed t o the combustor was a convergent-divergent exhaust nozzle. Two different ex- haust nozzles were used during t h i s investigation. The f l i g h t nozzle ( f i g , 2) was 42.2 inches long with a throat diameter of 38.8 inches and a 48-inch e x i t diameter (nozzle -throat t o combustor-area r a t i o , 0.66) .
This nozzle was used during t h e e n t i r e engine investigation, except the performance investigation when a heavy-duty convergent-divergent nozzle t h a t contained total-pressure rakes w a s i n s t a l l e d . The heavy-duty noz- z l e had t h e same dimensions a s the f l i g h t nozzle and thus had the same nozzle-throat t o combustor-area r a t i o .
The flameholder system is i l l u s t r a t e d i n figure 2 by a schematic diagram and i n figure 3 by a photograph.
The flameholder configura- t i o n consisted of a center p i l o t and three annular V-gutters intercon- nected by slanted r a d i a l V-gutters. The f u e l system consisted of the p i l o t f u e l and main f u e l system, The both using a common f u e l supply.
p i l o t burner contained three variable-area-type f u e l nozzles. The main f u e l system had three f u e l rings of 13-, 20-, and 26-inch diameter f i t t e d with LO, 15, and 40 variable-area-type fuel-spray nozzles, respectively.
The nozzles were rated a t 35 gallons per hour a t a d i f f e r e n t i a l pressure of 100 pounds per square inch. A l l the f u e l was sprayed downstream.
The f u e l used during t h i s investigation was MIL-F-5624B, grade JP-5, which has a lower heating value of 18,625 Btu per pound and a hydrogen- carbon r a t i o of 0.159.
Various methods of mounting the cooling l i n e r in the engine were The geometry of a l l the l i n e r s was the same (fig. 2 ) .
investigated.
The average cooling passage height w a s 0.78 inch except at the l i n e r i n l e t where the l i n e r was s l i g h t l y f l a r e d t o increase the cooling pas- sage area. The cooling l i n e r extended from 2.4 inches downstream of t h e engine diffuser i n l e t t o 10 inches downstream of the combustor o u t l e t .
A l l l i n e r s were made of 0.030-inch s t a i n l e s s s t e e l .
Engine s t a r t i n g t e s t s were made with f l a r e s i n s t a l l e d along the r a d i a l interconnecting V-gutters in the cambustor. Flares were elec- t r i c a l l y ignited with the f l a r e s pointing upstream i n some cases and downstream in others. Special fuels, which included boron t r i e t h y l and aluminum trimethyl, were used a s a source of spontaneous combustion in The boron t r i e t h y l had a specific gravity order t o s t a r t the engine.
Two types of aluminum trimethyl were used, one having a of 0.69.
specific gravity of 0,85 and t h e other a specific gravity of 1.3.
Gen- e r a l l y , about 1 t o 6 pounds of special h e 1 was injected through one o r more of the flameholder support pins a s i l l u s t r a t e d schematically in figure 2 and photographically i n figure 3.
The points of injection a r e a l s o indicated in figure 3.
Instrumentation Location of the engine instrumentation and amount of instrumenta- t i o n a t each s t a t i o n a r e shown in figure l. The f u e l flow was measured by a variable-area o r i f i c e .
Thermocouples (not shown) were placed on the engine shell, l i n e r supports, in the cooling-air passage, and on t h e cooling l i n e r . The l i n e r thermocouples were located a t four s t a t i o n s along the combustor with a t l e a s t three thermocouples placed a t t h e c i r - cumference of each s t a t i o n , Temperature readings were taken only t o indicate the order of metal temperatures and exact trends were not determined, PROCEDURE Flight Conditions and Engine Operation The investigation was conducted over a range of engine-inlet air flows from 60 t o 120 pounds per second, which correspond t o a l t i t u d e s of 73,000 t o 58,000 f e e t , respectively, a t a Mach number of 2.75 and at
i n l e t temperatures of 860°, 93s0, and 990' R . The range of f u e l - a i r
r a t i o s investigated f o r engine performance was from lean blow-out t o about 0.050. The cooling-liner durability w a s investigated a t a fuel- air r a t i o of approximately 0.045 and an i n l e t temperature of 990' R .
The engine diffuser operated s u p e r c r i t i c a l l y f o r a l l f u e l - a i r r a t i o s investigated .
The desired engine-inlet air flow was obtained by adjusting the t o t a l pressure and temperature a t the f r e e - j e t supersonic-nozzle i n l e t .
the f a c i l i t y exhaust The f r e e - j e t supersonic nozzle was f i r s t started, pressure w a s decreased u n t i l t h e engine nozzle was choked, the engine f u e l - a i r r a t i o was set, and then the engine was ignited by e i t h e r f l a r e s After ignition, the f u e l flow was varied over the o r special f u e l s .
The symbols and methods operable range t o obtain t h e performance data.
respectively.
of calculation a r e given in appendixes A and B, RESULTS AND DISCUSSION Engine Performance The performance of the f l i g h t engine f o r a range of i n l e t tempera- t u r e i s presented i n figures 4 t o 7 f o r engine-inlet air flows of 60, 80, 110, and 120 pounds per second. The performance data a r e a l s o pre- sented i n tabular form i n table I. The combustion efficiency, combustor- o u t l e t t o t a l pressure, combustor total-pressure r a t i o , and diffuser total-pressure recovery a r e shown a s a function of f u e l - a i r r a t i o .
I n general, a t a given a i r flow, t h e combustion efficiency increased while the combustor total-pressure r a t i o and diffuser total-pressure recovery decreased with an increase i n i n l e t temperature. A t approximately the design f u e l - a i r r a t i o of 0 "043 and an i n l e t temperature of 990° R, the combustion efficiency varied from 0.85 a t an air flow of 60 pounds per second t o about 0.88 a t an a i r flow of 110 pounds per'second. The cor- responding combustor total-pressure-ratio variation with a i r flow a t t h i s f u e l - a i r r a t i o was 0.795 t o 0.812. This performance compares closely with the performance of the heavy-duty version of t h i s engine as reported i n reference 2.
The f u e l - a i r r a t i o a t which lean blow-out occurred decreased ( f i g .
4 ) a s the i n l e t temperature was increased at a given a l t i t u d e , A t an a l t i t u d e of 73,000 f e e t , the fuel-air r a t i o a t which blow-out occurred decreased from 0.032 t o 0.028 a s i n l e t temperature was increased from 860' t o 990° R . There was only a s l i g h t e f f e c t of a l t i t u d e on lean blow-out as shown i n figure 8. A t an i n l e t temperature of 990' R, blow- out occurred between a fuel-air r a t i o of 0.027 t o 0.028 f o r the range of a l t i t u d e s investigated.
Engine Cooling The selection of a satisfactory method f o r attachment of the cool- ing l i n e r t o t h e burner s h e l l was the primary purpose of t h e l i n e r in- vestigation reported herein. The required cooling-liner length and cooling-passage height were determined i n the l i n e r t e s t s made during the previous investigation, which is reported i n reference 2.
The i n i t i a l l i n e r was made of s t a i n l e s s s t e e l (AISI 310) which was 0.030 inch thick. The method of l i n e r mounting, as provided by the manufacturer, was t o attach c l i p s t o both t h e engine s h e l l and l i n e r NACA FDI E55G22 with a Z-strip retainer.
The method of l i n e r attachment in t h e 30'- engine-diffuser section is shown here, while the mounting d e t a i l s of t h e ooling iner l i n e r in t h e combustor a r e i l l u s t r a t e d in t h e following sketch: The l i n e r was damaged ( f i g . 9) in t h e 30'-engine-diffuser section when a number of the Z-strips rotated out of the c l i p s a f t e r about 2 hours of cold-flow engine operation at an a l t i t u d e of 58,000 f e e t .
In an e f f o r t t o prevent the Z-strips from rotating out, new c l i p s were in- s t a l l e d on t h e cooling l i n e r i n the 30' conical engine diffuser beside t h e old clips, and new Z-strips, which conformed t o those installed i n t h e combustor section, were inserted. Additional cold-flow operation, however, resulted i n t h e opening of the l i n e r leading edge, which was repaired by wrapping and welding sheet metal around the leading edge as i l l u s t r a t e d i n t h e following sketch: During succeeding operation, differences in the thermal expansion of t h e c l i p s and l i n e r resulted i n ripples forming i n the cooling l i n e r a f t e r approximately 112 hour of burning time a t an a l t i t u d e of 73,000 f e e t .
This damage i s shown i n figure 10. In a few places the l i n e r pulled awqy from the engine s h e l l . The l i n e r was damaged extensively when an additional 33-minute burning run was made a t an a l t i t u d e of 58,000 f e e t - During t h i s run, hot spots were noticeable on the burner s h e l l . Maxi- mum engine s h e l l temperatures of 900' R and a difference of 600° R be- tween the s h e l l and cooling-liner metal temperature were measured, The l i n e r became detached from t h e burner s h e l l completely around t h e down- stream end and most of t h e spot welds holding the c l i p s pulled out.
The condition of the l i n e r i s shown in figure 1 1 a f t e r a t o t a l of 1 hour and 33 minutes of burning t i m e . Slight bulges also occurred in t h e ex- haust nozzle and engine s h e l l .
The damaged l i n e r was removed and a completely new l i n e r and com- bustor s h e l l were installed. This second l i n e r was modified in the 30'- engine-diffuser section with the new type of Z-strips previously des- cribed. In addition, 2;-inch L-strips were welded t o the c l i p s and l i n e r f o r added support. Small "hat" sections were spot welded t o t h e l i n e r , stradling the corrugations throughout the downstream half of the l i n e r . These modifications a r e shown by the following schematic dia- gram and a l s o by figure 1 2 .
The leading edge of the second l i n e r , which was formed by over- lapping the l i n e r metal, opened up as with the e a r l i e r l i n e r a f t e r 22 minutes of burning operation at an a l t i t u d e of 73,000 f e e t ( f i g . 13).
This damage was again repaired by wrapping and welding sheet metal around t h e leading edge.
With an additional 2-hour burning run, small ripples were present i n t h e l i n e r as shown i n figure 14, and i n addi- tion, hot streaks were observed on the engine s h e l l a s much a s 20 inches upstream of the exhaust-nozzle i n l e t during burning. A hole w a s burned i n the l i n e r when an additional run was made a t an a l t i t u d e of 58,000 f e e t f o r 69 minutes, making a t o t a l of 3 hours and 31 minutes of burning operation with t h i s cooling l i n e r . This f a i l u r e is shown i n figure 15, It was noted t h a t many spot welds a l s o f a i l e d throughout the l i n e r .
Another l i n e r of the same design as t h e second w a s i n s t a l l e d and t h e engine s h e l l was covered with an aluminum-foil blanket. A t o t a l of 1 hour and 43 minutes of burning time was accumulated with operation a t a l t i t u d e s from 58,000 t o 73,000 f e e t when f a i l u r e ocsurred ( f i g . 16).
Maximum blanket outside-surface temgeratwres of 1000 F were obtained.
Ln an attempt t o arrive at a more r e l i a b l e cooling-liner design, t h e combustion-chamber and exhaust-nozzle portion of the l i n e r was re- placed by t h e l i n e r configuration shown i n figure 17 and the following sketch: i o n s The l i n e r was made of s t a i n l e s s s t e e l that was 0.030 inch thick. The l i n e r was corrugated with h a t sections spot welded on t h e c r e s t s of the corrugations. Similar hat sections were placed on the engine s h e l l .
The l i n e r was inserted i n t o t h e engine s h e l l and rods were placed through To simplify the modifi- t h e hat sections t o r e t a i n t h e l i n e r in place.
cation, t h e section of t h e l i n e r in t h e 30°-engine-diffuser section was not altered. During operation hot streaks were seen along t h e unblan- keted engine s h e l l . Maximum engine-shell temperatures of about 1100~ F were recorded. After approxbnately 3 hours of operation a t a l t i t u d e s
of 60,000 and 67,000 f e e t , only a few small cracks developed 5 the
t r a n s i - l i n e r traiPing edge. The unmodified section of t h e l i n e r (30 t i o n section) f a i l e d as shown i n figure 18, and a f t e r an additional nwl qf 35 mbutes at an a l t i t u d e of 73,000 feet, t h i s t r a n s i t i o n section zmpletely buckled , The portion of t h e cooling l i n e r i n the combustor was not damaged and there was no sipg of deterioration o r f a i l u r e i n the modified por- t i o n of the l i n e r , This method of mounting the cooling l i n e r resulted i n approximately 35 hours of burning operation, although t h e l i n e r could have been used longer if t h e unmodified 3 0 ~ - e n ~ i n e - d i f f u s e r sec- t i o n had not f a i l e d .
Therefore, t h i s l i n e r design was believed t o be t h e most r e l i a b l e of those investigated.
Engine Ignition During cold-flow operation a t i n l e t conditions corresponding t o a l t i t u d e s of 58,000 t o 73,000 f e e t and with an unthrottled exhaust- nozzle area, it was found during as e a r l i e r portion of the program ( r e f . 2) t h a t ignition w a s impossible with spark plugs because of the low pressure and high velocities existing in the combustor. The use of f l a r e s and special fuels were thus adopted as alternative methods f o r ignition. The s t a r t i n g t e s t s reported herein were made at an i n l e t temperature of 410' F. This temperature was selected because it cor- responds t o the most c r i t i c a l condition f o r ignition, the cold day t e m - perature a t the s i t e where the missile is t o be f l i g h t t e s t e d .
The f l a r e s , varying i n number from four t o ten, were i n s t a l l e d along the r a d i a l interconnecting flameholder gutters i n t h e combustor t o provide t h e ignition source. A summary of the s t a r t i n g attempts made with f l a r e s is given i n t a b l e 13. The engine-inlet conditions and t h e number, position, and type of f l a r e s used f o r each s t a r t i n g attempt a r e shown in t h i s table. The chemical composition of the three types of f l a r e t h a t were used i s given in t a b l e 111.
The f l a r e s were considered unreliable as a source of ignition. O f the t h i r t e e n s t a r t i n g attempts t h a t were made with the f l a r e s , in f i v e cases a l l f l a r e s did not ignite, and i n only f i v e cases d i d the combus- t o r ignite. These r e s u l t s substantiate e a r l i e r flare-starting exper- ience during t h e heavy-duty engine investigation and in subsequent ex- perience with f l a r e s during flight-weight engine t e s t s .
The various s t a r t i n g a t t q t s t h a t were made with special fuels, boron t r i e t h y l and aluminum trimethyl, a r e summarized i n t a b l e ZV. This t a b l e includes t h e engine conditions, r a t e of f u e l injection, number of injection points, f u e l specific gravity, and whether ignition was suc- O f two s t a r t i n g attempts with boron t r i e t h y l , both were cessful o r not.
The engine was ignited in a l l cases with aluminum tri- unsuccessful.
methyl when t h e r a t e of f u e l injection w a s greater than about 2000 pounds per hour f o r the f u e l having a specific gravity of 0.85, and 3900 pounds The t o t d quantity per hour f o r f u e l having a specific gravity of 1.3.
The r e s u l t s with aluminum of f u e l injected varied from 1 t o 6 pounds.
trimethyl havfig a specific gravity of 0.85 were further substantiated by a subsequent investigation with an identical flight-weight engine.
Because t h e most experience by f a r was obtained with aluminum trimethyl having a specific gravity of 0,85, the r e s u l t s with t h i s f u e l a r e much more r e l i a b l e than those f o r t h e f u e l having a specific gravity of 1.3, Additional t e s t s a r e desirable t o further substantiate the s t a r t i n g r e - sults with aluminum trimethyl having a specific gravity of 1 , 3 .
S W Y O F RESULTS An investigation was conducted i n an a l t i t u d e t e s t chamber t o determine t h e performance, engine-burner-cooling, liner-design, and ignition characteristics of t h e XRJ47-W-5 48-inch-diameter ram-jet flight-weight engine.
The engine combustion efficiency varied from 0.85 a t an air flow of 60 pounds per second t o about 0.88 at an air flow of 110 pounds per sec- ond at the approximate design f u e l - a i r r a t i o of 0.043 and an i n l e t tem-
perature of 990' R . The corresponding combustor t o t a l -pres sure - r a t i o
variation with air flow at t h i s fuel-air r a t i o was 0,795 t o 0.812. Com- bustor lean blow-out occurred a t a fuel-air-ratio range from 0.027 t o 0 *028 a t an i n l e t a i r temperature of 990' R.
Three l i n e r configurations differing i n t h e type of l i n e r corruga- t i o n s and l i n e r support were investigated. With the best configuration,
t h e combustor section of the l i n e r withstood approximately 3 hours of
burning operat ion without significant damage, Aluminum trfmethyl was successful as a source of ignition at a l l engine-inlet conditions investigated provided t h a t the r a t e of injection exceeded about 2000 pounds per hour and t h e specific gravity of f u e l The f l a r e s were considered unreliable as an ignition source.
was 0.85.
The two attempts made with boron t r i e t h y l were unsuccessful.
Lewis Flight Pmpuls ion Laboratory National Advisory Committee f o r Aeronautics Cleveland, Ohio, July 26, 1955 S r n O I d j The following symbols are used in this report: A area, sq ft exhaust-nozzle-area coefficient
c~
exhaust-nozzle-expansion coefficient fuel-air ratio acceleration due to gravity, 32.2 ft/sec2 total pressure, lb/sq ft abs gas constant, 53 -4 ft-lb/(lb) ( O R ) total temperature, ? R flow rate, lb/sec ratio of specific heats combustion efficiency Subscripts : a air b combustor f fuel g gas 0 free-jet-nozzle inlet 3 engine -diffuser outlet 5 exhaust-nozzle throat
A i r flow, - Throughout t h e investigation t h e engine diffuser was
operated supercritically; therefore, the r e l a t i o n between i n l e t condi- t i o n s and engine a i r flow was unique.
The equation, which was obtained by calibrating t h e engine a i r flow, i s
Fuel-air r a t i o . - The f u e l - a i r r a t i o was calculated d i r e c t l y from
the measured f u e l flow and a i r flow:
Combustion efficiency, - The combustor temperature r a t i o was f i r s t
determined from t h e t o t a l pressure a t the exhaust-nozzle-throat area and i n l e t t o t a l pressure. The choked-nozzle equation and continuity equation is u But Wa = , and therefore,
-4%
where was determined from cold-flow t e s t s and a value of 0.987 was CA
used. The value of % was obtained from t h e exhaust-nozzle tempera-
t u r e and t h e expansion coefficient of Inconel. A trial-and-error method
was used i n determining r f o r the f i n a l T5 and appropriate f u e l - a i r
NACA RM E55G22 r a t i o . Ideal f u e l - a i r r a t i o was determined from ideal temperature-rise
curves and t h e value of 4- and i n l e t temperature. Combustion
efficiency was then calculated as the r a t i o of ideal t o actual f u e l - a i r r a t i o .
1. Povolny, John H., Farley, John M., and Hurrell, H. G . : Direct-Connect Investigation of Altitude Performance of the XRJ47-W-5 Ram- J e t - Engine Combustor. NACA RM E53J05, 1954.
2 . Welna, Henry J., and Smith, Ivan D . : Free-Jet Investigation of A l t i - tude Performance of XRJ47-W-5 Ram-Jet-Engine Conibustor, MCA RM E55C25, 1955.
3. Seashore, F e r r i s L., and Hurrell, Herbert G . : Starting and Perform- ance Characteristics of a Large Asymmetric Supersonic Free-Jet F a c i l i t y . NACA RM E54A19, 1954.
4. Farley, John M., and Seashore, F e r r i s L. : Full-Scale, Free- J e t Investigation of Methods of Improving Outlet Flow Distribution in a Side-Inlet Supersonic Diffuser. NACA RM E54L31aJ 1955.
TABLE I. - PRELIMINARY PERFORMANCE OF XRJ47-W-5 RAM-JET FLIGHT ENGINE AT MACH NUMBER 2.75 Combus- Fuel- Fuel Combustor Diffuser Exhaust- Engine Diffuser- Rbn Engine- Free- Free- air flow, tion total total nozzle total- outlet inlet stream stream ratio, W pressure pressure total pressure total air total total Wf, ciency, effi- ratio, recovery, pressure, ratio, pressure, flow, temper- pressure, lb/sec f ature, Pg/P3 nb P3/P0 P5 J P5/P0 Wa,O, 'a,~ P39 Po, lb lb lb/sec lb sq ft abs
sq ft abs '!'
sq ft abs Altitude, 58,000 ft ----- ------ 0.455 0.224 0.492 0 ------ ---- - .466 .231 -496 0 ------ ----- .462 .229 -495 0 ------ ----- .230 .496 0 819 .463 0.0298 0.846 .521 .407 .782 3.57 .0342 .854 .428 .803 4.11 1541 .534 .865 .542 .437 .806 4.31 .0364 .855 .552 .452 .819 4.79 .0399 .0495 .840 .591 .485 .821 5.98 .875 .507 .398 .780 3.55 .0295 ,886 .534 .419 .786 4.13 .0342 .870 .542 .438 .808 4.74 .0393 .858 .453 .819 5.35 .0445 1704 .554 .845 .467 .817 5.92 .0489 1758 .572 .886 .549 .448 .816 5.36 .0448 Altitude, 60,000 ft ------ ----- 0 0.2208 0.475 0.4648 1579 3397 1579 862 113.3 ---- - - - - - - - 0 .2259 .491 .4597 1618 17 3520 1618 992 109.4 0.0293 3.22 0.894 .392 .509 .771 1799 18 3537 1387 992 109.9 .894 .0341 3.75 .412 .782 .527 1865 19 3540 1458 991 110.1 .0393 4.33 .883 .430 .802 .536 1900 20 3544 1523 993 110.1 .0449 4.93 .875 .818 .447 .547 1935 3540 1583 109.8 21 996 .0496 5.44 .857 .458 .816 .561 1982 22 3531 1618 994 109.6 Altitude, 67,000 ft ------ ----- 0 0.230 0.507 547 0.453 79.2 2379 1078 865 ------ ----- 0 .22 9 .494 586 .464 79.2 2554 24 1186 998 0.0293 2.34 0.788 .769 .396 947 .515 79.8 2392 1231 25 861 .0342 2.74 .854 .795 .429 1026 .538 80.1 2397 26 1290 859 .0390 3.12 .856 .812 .449 1075 .553 80.0 2394 1324 27 858 .0439 3.53 .845 .813 .466 1119 .574 2400 28 1377 80.5 852 .0487 3.92 .831 .818 .480 1153 .586 80.5 2403 1409 854 .0288 2.31 .854 .394 .765 981 .515 80.3 2492 1283 30 923 .0343 2.76 .880 .421 .791 1051 .533 80.4 2495 1329 923 .0390 3.13 .862 ,795 .436 1091 .549 80.3 2501 1372 929 .0440 3.52 .859 .809 .452 1130 .559 80.0 2500 1397 33 935 .0488 3.92 .830 .809 .463 1159 .573 2501 1432 80.3 34 930 .0296 2.36 .875 .764 .390 1002 .511 79.9 2568 1311 35 991 .0341 2.74 .871 .410 .773 1054 .530 80.3 2573 1363 984 .0392 3.14 .875 .791 .429 1103 .542 80.1 2573 1394 989 .0439 3.52 .852 .801 .441 1137 .551 80.2 2574 1419 38 988 .0491 3.93 .829 .807 .453 1166 .561 80.0 2574 1445 39 993 Altitude, 73,000 ft 0.828 0.0346 2.08 0.794 0.426 764 0.536 60.0 1795 962 40 857 .844 .0392 2.35 .815 .447 802 .548 59.9 1795 984 41 860 .866 .0397 2.36 .El5 .448 809 .550 59.5 993 1805 42 883 .841 .0427 2.60 .815 .461 837 .566 60.9 1027 1814 43 851 .838 .0438 2.62 .El0 .462 829 .570 59.8 1795 44 1023 864 .823 .0492 2.92 .821 .475 850 .578 1035 59.3 1790 45 874 .846 .0492 2.93 .826 .479 862 .580 1044 59.5 1800 46 877 .845 .0348 2.09 .785 .416 780 .529 60.1 1877 47 993 934 .865 .0384 2.31 .795 .434 813 .5+6 60.1 1874 48 1023 931 .840 .0432 2.61 .807 .449 844 .556 60.5 1880 49 1046 926 .828 .0489 2.93 .807 .463 865 .573 1072 59.9 1870 50 934 .855 .0345 2.07 .776 .408 786 .526 60.0 1926 51 1013 988 .856 .0388 2 : 3 4 .793 .424 .535 822 60.3 1938 52 1037 991 .844 .0429 2.59 .796 .437 848 .549 60.4 1940 53 1065 989 -833 .0490 2.93 .808 "452 .560 873 59.8 1930 54 1080 998
TABU 11. - S m Y OF ENGINE STARTING WITH FLARES
I n s t a l l a t i o n Engine Remarks S t a r t i n g S t a r t i n g c o n d i t i o n F l a r e s knited attempt Along A i r Number Along I n l e t D i r e c t i o n Fuel- Type i n n e r o u t e r a i r flow, i n s t a l l e d temper- of flame l b / s e c s l a n t s l a n t r a t i o , a t u r e , from g u t t e r s OF g u t t e r s f l a r e f / a No E i g h t f l a r e s i g n i t e d 5 1 120 40 1 0 5 Downstream 0.026 No 5 F i v e f l a r e s i g n i t e d 5 .045 2 40 1 0 410 Upstream A l l f l a r e s i g n i t e d , 23 s e c l a g 5 3 4 0 s 5 380 Downstream .045 1 0
-
8 .043 E i g h t f l a r e s i g n i t e d 4 8 400
-
8 E i g h t f l a r e s i g n i t e d 5 8 410 .032
-
Four f l a r e s i g n i t e d 6 8 410 8 .030
-
o 4 A l l f l a r e s i g n i t e d 7 4 405 .042
-
8 S i x i g n i t e d 8 8
-
A l l f l a r e s i g n i t e d 30 110 8 9 8
-
8 A l l flames i g n i t e d 1 0 8 3 0 110
-
s 8 A l l f l a r e s i g n i t e d , 3 s e c l a g 11 8
-
8 A l l f l a r e s i g n i t e d 1 2 8 No
-
No 8 Seven f l a r e s i g n i t e d 13 8
TABLE 111. - FLARE COMPOSITION
Chemical composition of f l a r e s , p e r c e n t by weight TABLE IV. - SUMMARY OF ENGINE STAXTING WITH SPECIAL FUELS Jet &if fuser , Periscope --7 I I Plan view I I a
Instrumentat ion i I
station 0 3 5 ~p Wall static-pressure Static-pressure Thermocouples Instrument Total-pressure orif ices tubes station tubes 0 14 21 48 21 0 0 5 32 Figure 1. - Free-jet installation of 48-inch ram-jet engine.
I - ' Figure 2 . - Cross section of XRJ47-W-5 ram- jet flight engine. ( ~ l l dimensions in inches .)
Figure 3. - Special fuel-injection manifold with four injection points.
Fuel-air r a t i o , f/a Figure 4.. Engine performance. Altitude, 73,000 f e e t ; i n l e t a i r flow, 60 pounds per second.
-4 -4 2 ,n -80 0 F let t o t a l temperature, To, O R .rl Q 1 " D ti 8 -70 Fuel-air r a t i o , f / a Engine performance. Altitude, 67,000 feet; i n l e t a i r Figure 5.
flow, 80 pounds per second.
A = .
d 0 4 .rl O c , - P a k & 0 a,
% 2
5 rn .a m
g c
u a Fuel-air r a t i o , f / a Figure 6 . Engine performance. Altitude, 60,000 f e e t ; i n l e t a i r flow, U 0 pounds per second; i n l e t temperature, 990°R.
Fuel-air r a t i o , f/a Figure 7. Engine performance. Altitude, 58,000 feet; i n l e t a i r flow, 120 pounds per second.
Fuel-air r a t i o , f/a Figure 8. Znginer performance. I n l e t t e m p r a t u x e , 9 9 0 ~ ~ .
Figure 9. - Damage to 3 0 ' conical transition section.after 2 hours of cold-flow operation.
Figure 10. - Damage to original liner after l-hour of burning operation at altitude of 73,000 feet.
( a ) Complete engine.
Figure 11. - Failure of original l i n e r a f t e r a t o t a l burning time of 1 hour and 33 minutes.
(b) Exhaust nozzle removed.
Figure 11. - Concluded.
Failure of original l i n e r a f t e r a t o t a l burning time of 1 hour and 33 mintues.
Figure 12. - Modification of second cooling liner.
Figure 13. - Leading-edge dam~ge of second cooling l i n e r a f t e r a t o t a l burning time of 22 minutes.
Figure 14. - Condition of second cooling l i n e r a f t e r a t o t a l burning time of 2 hours and 22 minutes.
(a) Hole burned i n liner.
Figure 15. - Damage t o second cooling l i n e r a f t e r a t o t a l burning time of about 3$ hours.
(b) Spot-weld f a i l u r e .
Damage t o second cooling l i n e r a f t e r a t o t a l burning time of 32 hours.
Figure 15. - Concluded.
Figure 16. - Failure of t h i r d cooling l i n e r a f t e r a t o t a l burning time of 1 hour and 43 mimtes.
(a) Cooling liner.
Figure 17. - NACA cooling-liner design.
NACA RM E55G22 (b) Engine shell.
Figure 17. - Concluded. NACA cooling-liner design.
Figure 18. - Damage to 3 0 ' conical transition section with NACA liner installation.
NACA R M E55G22 Aeronautical Research Fac ilit ies W i 1 l b . m A. Fleming Aeronautical Eesearch Sc f a t f st Propuls ion System Bruce T. L u n d b Chief Engine Research Division