Document
Restriction/Classification Cancelled Restriction/Classification Cancelled ?^'^ f ^ ^^'^^
U
NACA MI No. E7A20a NATIONAL ADVISORY COMMITTEE FOR AER.^ALTDAUTICS RESEARCH tUMORA1TM1 f or the Air Materiel Command, Army Air Forces FLIGHT INVESTIGATION OF THE EDE CTS OF ICE ON AN.I-16 JET-PROPULSION ENGINE Pragliola and Milton Werner By Philip C.
SUMARY A flight investigation of an 1 16 : jet-propulsion engine installed in the waist compartment- of a B-24M air- .-lane was made to determine the effect of induction-system icing on the performance of the engine.
Flights were made a t inlet-air temperatures of 15 0 ) 20 0 , and 25 0 F, an of indicated airspeed 1.80 miles per hour, jet-engine speeds of 13,000 rpn^ liquid-water contents of aplroxiLlately 0.3 to 0.5 gram, and 15,000 per cubic meter, and.an average water droplet size of approximately 50 microns.
Under the most severe icing conditions obtained, ice formed on
the screen over the frcnt inlet to the compressor and obstructed about
70 percent of thk , front-inlet area. The thrust was thereby reduced
13.5 percent , , the specific fuel_ consumption increased 17 percent, and
the tail-pipe temperature increased 82 0 F. No icing of the rear-
compressor-inlet screen was encountered.
INTRODUCTION 19ACA Cleveland A flight investigation has been conducted at the laboratory to de te. , L^nine the effect of induction-system icing on the performance of an 1-16 jet- p ropul r , , ion engine. The engine alone was investigated and no attempt was made to determine the effect of icing of the inlet duct to the engine. The engine was located in the waist compartment of a B-201 airplane.
Data are presented to show the — effect of icing of the induction system on jet thrust ., engine speed, tail pipe temperature, and n
P ro I I Al
Ft
PfX L
NACA M4 No. E7A20a s-:Decific fuel consum.otion. These data are compared ;rith results of flights made when -the front inlet of the engine was obstructed by alLuiinum. bands.
INSTALLATION The I-16 jet-proDulsion engine has a twin-inlet centrifugal compressor and is rated at 1600 pounds of thrust at an engine speed of 16,500 rpm, The engine was installed in the waist compartment of the B-24M airplane, as shown in figures 1 and 2. In order to facili- tate the investiGation, all turrets were removed from the airplane and replaced by suitable fairing. The leading edges of the wing (except inboard) and tail surfaces were thermally de-iced by four exhaust-heat exchanClers. An induction-air inlet was installed on top of the airplane to 12rovide induction air to the engine (fig. 2).
Windows were installed in this duct in order to observe the f orma- tion of ice on the front and rear air inlets to the compressor and in the inlet duct. Exhaust gas was ex)elled through a long tail pipe that was provided with external cooling.by air sup;,.Dlied from two scoops, one located in each waist ^;unner l s window. Air at a pressure of 100 ponds per square inch and , water at a pressure of 10 pounds per square inch were externally mixed by t3he use of nozzles that were installed J feet in front of the induction-air inlet, This water-air spray, which was directed towards the inlet, was used to simulate icing conditions and also to raise the liquid-water content of the air at the engine inlet during natural icing conditions.
INSTRUMENTATION The engine was instrumented as shown in figure 3. In addition to standard measurem.ents of engine speed, burner-fuel pressure, oil pressure, tail-piy)e temperature, and bearin temperatures, the following measurements were obtained.
(1) Surveys of the inlet air were made by iron-constantan thermocouples unshielded from ice and total-pressure tubes located in both front and fear inlets to the compressor. The pressure tubes, in the compressor inlets were heated -to prevent icing.
(2) Free-strear_: static pressure was measured by flush orifices located on the sides of the airplane.
(3) Static pressure at the inlet to the compressor zras measured by a flush orifice located in the duct over the front compressor inlet.
NACA RM No. E7A20a (4) Chromel-alumel thermocouples were spot-welded to the out- side of all the burner-outlet elbows to indicate burner combustion.
(5) A calibrated survey ring was installed in the tail pipe to measure tail-pipe temperatures and pressures for the determination of jet thrust. These thermocouples also were of the chromel-aluxel type.
(6) Ice formation on the front inlet to the compressor was photographed by a 35-millimeter moving-picture camera mounted next to the jet engine.
The liquid-water content of the air entering the engine was determined by means of a rotating cylinder ., one-eighth inch in diameter (fig. 4) located just ahead of the front compressor inlet.
(fig. 3). This cylinder was inserted about once every 8 minutes during icing runs for a 5-minute period. From the weight of ice collected ., the average diameters were determined. The liquid-water content of the air was then directly determined from the volume of air intercepted and the weight of ice collected, inasmuch as the collection efficiency of small-diameter cylinders (one-eighth in.)
approaches 100 percent (reference 1).
liater-droplet size for simulated icing runs was determined by means of a droplet sampler ., as shown in figure 5. Use of this
droplet sampler and method of analyzing the slides is given In
detail in reference 1. By use of a sooted slide, a droplet sample could be taken in flight and analyzed later. The slide is exposed for 0.01 second and droplets collect on the sooted surface. As each droplet hits the sooted surface, it flattens because of impact pressure. The slide is treated with a nonwetting agent before it is sooted; thus, when the droplet returns to a partial sphere after the force of impact is lost, the soot clings to the water rather than to the slide. After the water evaporates, a black spot equal to the diameter of this partial sphere is left on the slide. This diameter is divided by a factor of 1.25 to determine the true droplet diameter.
Iron-constantan thermocouples were accurate to ±2 0 F and chromel- alumel thermocouples to ±7 0 F. Calibration data of gages used for measuring tail-pipe pressures deviated from the average curve by a maximum of 0.1 inch of mercury. The measurements of liquid-water e content and droplet size were made as accurat ly as current instru- ments permit; however, the state of this method does not permit a of definite statv.jent accuracy.
NACA RPi No. E7A20a CONDITIONS AND PROCEDURE Flights were made at three altitudes in order to obtain front inlet-air temperatures of 15 0 , 200 , and 250 ; at each altitude an indicated airspeed of 180 miles per hour was maintained. All perform- ance data presented, however, are reduced to standard sea-level condi- tions to eliminate variations due to different air densities. The engine was operated at speeds of 13,000 and 15,000 rpm corresponding to low and high normal-cruise conditions.
Simulated icing conditions and a combination of natural and simulated engine operating conditions were used. Only light natural- icing conditions were encountered, but it was possible to augment the natural icing with water sprays that produced liquid water at the rate of approximately 0.3 to 0.5 gram per cubic meter with an average droplet size of 50 microns at the front inlet to the compressor. The most severe icing conditions were obtained in flight 3. Greater water contents similar to severe natural-icing conditions were desirable but were unobtainable with the equipment installed in the airplane.
Additional runs were made with the front inlet to the compressor obstructed by alu-minuet bands.
SY1VM OLS The following symbols are used in this report: A tail-pipe exit area, 0.818 (sq ft) Ar tail-pipe area at survey ring, 1.108 (sq ft) Cl; C% constants Fj jet thrust, (lb) Fn net thrust, (lb) f specific fuel consumption, (lb fuel/hr/lb thrust) g acceleration of gravity, 32.2 (ft/sec2) H total pressure, (lb/sq ft absolute) k calibration factor N engine speed, (rim.)
NACA RM No. E7A20a p static pressure, (lb/sq , ft absolute) gas constant ((ft-lb)/(slug)(oF)) R T temperature, (OR) V true airspeed, (ft/sec) air flow,
Wa (lb/sec)
W f fuel flow, (lb/hr)
W g gas flow, (lb/sec or lb/hr)
ratio of specific heats, 1.33
Y Subscripts: corr corrected -to standard sea-1--vel conditions r tail-pipe survey ring 0 ambient free stream METHOD OF CALCULATION The thrust was calculated from temperatures and pressures obtained with the survey ring located in the tail pipe.
The jet thrust F J was obtained from 44
y
r ►
- £ C1p0 - ^ j j ^,pu_))
l f
where C 1 - 'Y-1 Ak The net thrust Fn was found by subtracting the initial momentum of the inlet air from the jet thrust.
WaVO
Fn=Fj - ^
6 NACA RM No, E7A20a 01; The gas flow W was obtained from the equation 9- Y_l (Hr) 7 - I + 0.6 119 = C2PrPr) Tr where C = IcArI y-1 V - 1 TR The air flow is then found by subtracting the fuel flow from gas flow Wa = % - Wf The net specific fuel consumption is given as Wf/Fn- All performance data were corrected to standard sea-level condi- tions at the compressor inlet by the following correction factors: compressor-inlet total pressure H, (lb/sq ft absolute standard sea-level pressure (2117 lb/sq ft absolute) 0QN compressor-inlet t emperature T ( - / 2M e standard sea-level temperature (518 0 R) The following equations show the method of correcting the various performance variables: Fn F n, corr - 8 N Neorr - -
'if BB
= Wg W 9, corr Wa Vj Ila., corr - NACA RM No. E7A20a 7 Wf , e Wf, corr 2-- S Tr Tr, corr = 8 fcorr - f RESULTS AND DISCUSSION Data taken during three flights (1,2, and 3) in which natural- and simulated-icing conditions were combined are shown graphically in figure 6. The conditions and results are presented in table I.
From this table it can be seen that the loss in thrust varied from 5.5 to 13.5 percent, the tail-pipe temperature increased from 270 to 900 F, the specific fuel consumption increased from, 6.5 to 17.0 percent, and the percent obstruction varied from 32 to 75 per- cent. Natural-icing conditions encountered during all three flights resulted in a band of ice approximately 6 inches wide and about 1/2 inch thick on the leading edge of the unprotected inboard section of the wing.
Ice formations on the screen over the front inlet to the compressor during the three flights are shown in figure 7. These photographs were taken when the maximum change in performance was noted.
During flight 2, the nozzles froze after 27 minutes (fig. 6) and the ice on the inlet sublimated. Some ice collected in the bends of the duct and at the duct inlet but it is believed that the effect of this ice on the performance of the jet engine has no significance on the results reported.
The air that entered the rear inlet was required to pass the compressor-outlet ducts (fig. 3), which heated the inlet air sufficiently to maintain above freezing temperatures at the rear screen during all three icing flights and therefore prevented ice formation on the rear screen. At lower ambient-air temperatures, the temperature of the rear screen might be below the freezing level. In this case, if very small droplets would reach the screen without separating from. the air stream at sharp bends, ice would result.
8 NAOA RM No. E7A20a Aluminum.bands were placed over the front inlet to the compressor (fig. 8) to determine the change in performance with 25, 50, and 75 percent of the front-inlet area obstructed, The bands were located to the rear of the inlet to represent the ice formed in the previous left unrestricted because no ice formed flights. The rear inlet was in this inlet. Results of flights with the restriction bands are shown in figure 9. From these data, it is possible to determine the change in performance when a certain percentage of the front inlet is obstructed. The maximum loss in net thrust with 75 percent of the front inlet obstructed was about 13.5 percent at 13,000 rpm and 31 percent at 16,500 rpm. The increase in tail-pipe temperatures at these speeds and 74-percent obstruction of the front compressor inlet was 82 0 F at 13,000 rpm and 166 0 F at 16,500 rpm. Specific fuel consumption increased 24 percent at 13,000 rpm. and 35 percent at 16,500 rpm.
The tailed points on figure 9 are the maximum changes in performance of flights 1, 2, and 3 taken from figure 6. An approxi- mation of the amount of blocking of the front inlet from the ice when the average droplet size was 50 microns can be made from these points.
In the case of flight 3, with a water content of 0.4 to 0.5 gram per cubic meter and a front inlet-air temperature of 25 0 F. the inlet to the compressor was obstructed about 70 percent. With a water content of 0.3 to 0.4 gram per cubic meter and a front inlet-air temperature of 15 0 F the obstruction for flight 2 was about 60 percent. For flight I with an inlet-air temperature of 20 0 F, the inlet area was reduced somewhere from 30 to 50 percent. These approximations of the obstruction of the front inlet based on changes of engine performance satisfactorily agree with comparison of photographs of the ice in figure 7 and the artificially obstructed inlet shown in figure 8.
Attempts in clear air to simulate natural-icing conditions and to cause blocking of the compressor inlet were unsuccessful because of the large size of water droplets and unsaturated atmospheric air.
Best results were obtained when water sprays were used in conjunc- tion with natural ice.
After 54 hours of running time, the I-16 engine still operated satisfactorily. During about 20 percent of this time, the front inlet was partly blocked by either ice or by the obstructing bands.
These conditions had no noticeable permanent effect on the rotor bearings or on the engine itself. During runs when the inlet was blocked, there were no indications of poor combustion or rough engine operation.
NACA RK No. E7A20a SUMMARY OF RESULTS The following results were obtained from the icing investiga- tion of the 1-16 jet-propulsion engine installed in the waist compartment of a B-24M airplane: 1. With an average liquid-water content of 0.4 to 0.5 gram per cubic meter, an average droplet size of 50 microns, and a front of inlet-air temperature 250 F, ice formed on the screen over the front inlet to the compressor and blocked the front inlet approxi- mately 70 percent. The obstruction of the front inlet to the compressor produced by the ice formation reduced the thrust 13.5 percent ) increased the specific fuel consumption 17 percent, and increased the tail-pipe temperature 82 0 F.
2. No icing of the rear inlet screen occurred because the inlet-air temperature of the rear inlet to the compressor was above freezing in all cases. With lower temperatures and small water droplets it is possible that some ice would build up on this screen.
Aircraft Engine Research Laboratory, National Advisory Conaittee for Aeronautics, Cleveland . , Ohio.
Philip C. Pragliola, Aeronautical Engineer.
Milton Werner., Mechanical Engineer.
Approved: Lewis A. Rodert, Aeronautical Engineer.
Abe Silverstein; Aeronautical Engineer.
jh NACA RM No. E7A20a REFERENCE 1.
Vonnegut, B^, Cunningham.,, R. M.., and Katz ., R.
E.: instruments for Measuring Atmospheric Factors Related tc Ice Formation on Airplanes. Dept. Meteorology, De-Icing Res.
Lab., M.I.T., April 1946.
NACA RM No. E7A20a TABLE I - CONDITIONS AND RESULTS OF ICITtiG FLIGHTS I ?
Flight CONDITIONS _f Pressure altitude, ft 11'7,000 19,500 5000 _ Inlet-air temperature, of Front 20 15 25 Rear 36 33 35 180 180 180 Indicated airspeed, mph 13,000 Jet--engine speed, rpm 15,000 15,000 Liquid-water content at 0s3 to 10.4 to ! front compressor inlet, 0.4 gram/cu meter (a) 0.5 Average droplet size, microns 50 50 50 Length of time in icing conditions, min ' 30 27 RESULTS Loss in thrust s i percent 5,5 10 13,5 pounds 50 100 80 percent obstruction c 49 61 75 i Increase in tail-pipe 27 90 82 temperature, oFb 32 60 7 4 !
percent obstruction c Increase in specific fuel consumptionb !
percent 6.5 9 17 lb/(hr)(lb thrust) 0111 0.15 ! 0.36 percent obstructionc 47 ! 53^ 6 8 i allot determined.
b Corrected to standard sea-level conditions.
c Taken from figure 9.
NATIONAL ADVISORY COM,1ITTEE FOR AERONAUTICS A n z M Figure la — B-244 airplane
with 1
-16 jet—propulsion engine installed in waist compart- ments J Z D C7 D Z O M J A N O A) ioI ing- ;oop sion is nel ro I s Total for a s N Figure 2, — B-24 fuselage showing location of jet engine and equipment for tests.
D n D Z1 z O m a D Pv W p up I e tubes vey COMMITTEE FOR AERONAUTICS cta Figure 3. - Location of temperature, pressure, and meteorological measuring instruments on 1-16 engine.
U4 wACA Am No. E7A20a ^
MW
' C- 15608 o ' 13- 46 Figure 4. — Rotating c j[/ lnder used for determination of wzter content.
.
—' NA Fig, 5 A Figure 5. — Droplet sampler used for determination of water— droplet size.
No. E7A20a rig. 6 NA CA RU ^ U N Sr- ¢^ 0 4D 'CS Cd a 2 U O ^a rd Gx.
P4O e-9 +i d Id S+ -P :3 ^a ^d a-1 La 9 a fd i' U H N d, 4i 41 U V i.
a-i 4) U ^ 4D ^ LL O ^4.3 10 r4 P4 to r. 0 C o td a U d-3 4 G ^ V j„ U fe ^' i 41 4) w Q 4^ .^ Se 0 g Time, min Figure 6. ® Engine performance during icing runs of flights 1, 2, and 3® All values corrected to standard sea-level conditions.
( a ) No i c e . ( b ) F l i g h t 1 , NACA C - 17457 12-31- 46 ( d ) F l i g h t 3.
( ca F l i g h t 2, n -.
F i g u r e 7 . - F r o n t c o m p r e s s o r - i n l e t s c r e e n w i t h o u t a n d w i t h i c e d u r i n g f l i g h t s 1 , 2, a a n d 3 .
Z D C7 D . '0 Z O m D N ro cn Figure g® — Location of aluminum bands used to block front compressor inlet.
Go NACA RM No. E7A20a ^ U N c~. ° 0 vl 4r zy F:J N M g2 U v a ^w ^ 4 ^0 ^Co a^ cd +) U r-1
P 40
^ 44 d U U Ls a 30 U a S'.
P, 0 20 q a 0 10 o fd U 4 0 U N N f' .. U 6". N +rl ¢.
d, a ^ U d^ Obstruction, percent Figure 9. - Engine performance during' flights with alumiim restriction bands on front compressor inlet and compari'.n of these data with data from icing flights. All values corrected to standard sea-level cond°.tions.