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*5EP 161946 M '

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TECHNICAL NOTE

No. 1109

FLIGHT INVESTIGATION OF THE COOLING CHARACTERISTICS

OF A TWO-RCW RADIAL ENGINE INSTALLATION

II - COOLING-AlR RRESSUKE RECOVERY AND ERESSÜRE DISTRIBUTION

By E. John Hill, Calvin C. Blackman, and James E. Morgan

Aircraft Engine Research Laboratory Cleveland, Ohio

N HP^ Washington July 1946

N A C A LIBRARY LANGLEY MEMORIAL ABRONAUTICAL

LABORATORY Lantjley Field, Vsu

433 3067

NATIONAL ADVISORY COMMITTEE .FOE AERONAUTICS

TECHNICAL NOTE NO. 1109

FLIGHT INVESTIGATION OF THE COOLING CHARACTERISTICS

OF A TWO-ROW RADIAL ENGINE INSTALLATION

II - COOLING-AIR PRESSURE RECOVERY AND PRESSURE DISTRIBUTION

By E. John Hill, Calvin C. Blackman, and James E. Morgan

SUMMARY

Flight tests have "been conducted at altitudes of 5000 and 20,000 feet to investigate the cooling-air pressure recovery and distribution for a two-row radial engine enclosed in a low-inlet- velocity cowling of a twin-engined airplane. The effect of flight variables on average recovery and circumferential, radial, and longitudinal distribution are presented for level flight; also included is a comparison of pressure-drop measurements across the engine, as indicated by nine different combinations of pressure tubes.

The results of these tests showed that pressure recovery and distribution can be greatly affected by changes in flight variables. Those variables having the greatest effect were cowl-flap angle, angle of attack of the thrust axis, and the propeller thrust disk- loading coefficient. The tests further showed that large differences, sometimes amounting to 100 percent, were obtained in the results indicated by various methods of measuring pressure drop across the engine. •-_

r

On the basis of the results, it is observed that an important consideration in the design of cowlings and cowl flaps should be the obtaining of good distribution of cooling air, as well as minimum drag for the installation. The fact that these tests showed that the front recovery decreased with an increase in propeller thrust disk- loading coefficient provides additional evidence that the recovery is greatly affected by the combined propeller-nacelle design. Also of significance is that a large increase in front recovery in these tests resulted in a similar increase in rear pressure, indicating that an 'increase in the front recovery of an air-cooled engine is not always an effective method of increasing the cooling-air flow.

NACA TN Wo. 1109

INTRODUCTION

A flight Investigation of a two-row radial engine enclosed in a low-inlet-velocity cowling was undertaken to determine the cooling characteristics of the installation at altitude. The introductory- report of this investigation (reference l) was concerned with the correlation of the engine cooling variables at altitude by the method-

described in reference 2 and the adaptability of this correlation for the determination of the general cooling performance of the engine installat Lon.

The present report is a study of the cooling-air pressure recovery and distribution within the engine cowling. The distribution of cooling-air flow is one of the important factors that control the distribution of temperature among the cylinders of an air-cooled multlcylinder engine. Inasmuch as efficient engine operation postu-- latos a relatively uniform temperature distribution in order to minimize cooling drag and to develop mn-rimwm power and fuel economy, a study of the factors controlling cooling-air distribution is of considerable importance.

The quantity of cooling air flowing over the Individual cylinders of an air-cooled engine is mainly a function of the pressure drop across the cylinder. This pressure drop is determined by the pressure recovery and distribution at the front and the rear of the engine, which In turn are dependent upon the cowling design, flight conditions, and engine conductivity (a nondimensional factor Indicative of the resistance to cooling-air flow through the engine). A large number of wind-tunnel and flight investigations have been made involving cooling- air pressure recovery and average pressure drop but they have been •associated mainly with the problems of optimum cowling design. Little work has been reported concerning the effect of flight variables on the distribution of cooling-air flow.

An investigation was made at the NACA Cleveland laboratory of the cooling-air pressure recovery and distribution throughout an air-cooled engine installation and of the effect of important flight variables on recovery and distribution during level unaccelerated flight. The results are, in detail, applicable only to this engine installation; however, in the discussion an attempt is made towards a general interpretation of the results, A study of uverage cooling- air pressure recoveries and circumferential, radial, and longitudinal pressure distribution is Included. The variables investigated wore: (a) airplane speed, which influences the pressure available for cooling the engine; (b) cowl-flap angle, which changes the resistance to air flow through the cowling and also effects the cowl-exit pressure; (c) angle of attack of the thrust axis, which influences tho characteristics of air flow Into the cowling; (d) propeller thrust

MCA TN No. 1109

disk-loading coefficient, which, is a measure of the pressure rise across the propeller; and (e) propeller speed, "which affects the rotation imparted to the air. A comparison of different pressure- drop measurements across the engine is also included.

SYMBOLS AMD COEFFICIENTS

The following symbols are used in the analysis of the investigation:

Cs speed-power coefficient, ip/pV5/P N2

D diameter of propeller

H total pressure above atmospheric static pressure

N propeller rotational speed

P power absorbed by propeller

p static pressure above atmospheric static pressure

Ap pressure drop across engine v.

q.c free-stream impact pressure 2

S propeller-disk area, ——-

T thrust, P n/V

T thrust disk-loading coefficient, T/a,S

V velocity relative to air stream

V/HD propeller advance-diameter ratio

oc angle of attack of thrust axis

ß blade angle of propeller at 0.75 radius

n propeller efficiency

p mass density of free stream

NACA TIT No. 1109

Subscripts

ae average engine

b barrel

e exhaust side of cylinder

fr front row

h head

i intake side of cylinder

rr rear row

t top of cylinder

1 to 8 logitudinal stations relative to cylinder

APPARATUS AND METHODS

Airplane and engine. - The investigation of cooling-air pressure recovery and distribution was conducted on the right engine installation of a twin-engined airplane (fig. l). A sketch of the cowling with charge-air and oil-cooler ducts is shown in. figure 2. The cowling is of the short-nose type without entrance diffuser and with cowl flaps located on both sides of the lower portion of .the nacelle. Very little exit area is provided for the cooling-air flow from the_ cowling except through the cowl flaps, which remain partly open even in the "full-closed" position. The test engine was of the 18-cylinder, double-row radial, air-cooled type having a gear-driven, single-stage, two-speed supercharger. The conventional propeller reduction gear, which had a ratio of 2:1, was replaced with a torquemeter having the same ratio.

The propeller was four-bladed, 13^ feet in diameter, and of the

constant-speed type; it was fitted with cuffs and spinner that are standard for this installation.

Approximate normal flight conditions for the airplane at a gross weight of 30,000 pounds are given in the following table for level flight at altitudes of 5000 and 20,000 feet for take-off and for climb:

NACA TN No. 1109

Operating condition

Altitude (ft)

Brake horsepower per engine

Engine speed (rpm)

Indicated airspeed

(mph)

Angle of attack of thrust axis

(dog)

Cowl- flap position

Low-power cruise

5,000 20,000

800 800

1800 1900

195 170

4 6

Closed Closed

Normal cruise

5,000 20,000

1050 1050

2100 2300

220 195

3 4

Closed Closed

Rated power

5,000 1500 2400 255 1.5 Closed

Climb 1250 2400 170 Open Take-off 1850 2600 110 Open

Instrumentation. - The relative location of all pressure tubes is shown in figure 3. The cooling-air pressure in front of the engine was measured by shielded totalspressure tubes on rakes (in front of four cylinders only), total-pressure tubes at the baffle entrance, and by tubes placed on the head baffle that butts_ against the sealing ring of the cowling. Pressures behind the engine were measured by open-end tubes in stagnant regions and by total-pressure and closed-end static-pressure tubes downstream from the cylinder. Copper tubing of 1/8-inch, diameter was used for all pressure tubes; the designation, type, and exact location of the pressure tubes are given in table I. The cylinder numbering system used in the table" and throughout the report is conventional; the cylinders are numbered clockwise when viewed from the rear of engine with cylinder 1 being the top rear-row cylinder.

A diagrammatic sketch of the system used for measuring the pressures is shown in figure 4. The pressures were recorded by NACA 30-cell and single-cell recording manometers and by a 100-tube liquid-manometer board photographed in flight. The 30-cell manometer consists of 30 differential-pressure cells in conjunction with seloc-_ tor valves and permitted an accurate recording of 254 pressures (including reference pressures) within 35 seconds. The 100-tube liquid-manometer board was connected to a two-bank, 100-tube selec- tor valve enabling the photographing of two consecutive sets of _ pressures. Of these 200 pressures, 12 were used to establish the _ reference line for the manometer board. The reference pressure for •both the 30-cell recording manometer and the 100-tube liquid manometer was the boundary-layer static pressure obtained by a flush orifice in the bottom of the fuselage.

The free-stream static pressure was measured by a calibrated swiveling static-pressure tube mounted on a boom extending 1 chord

NACA TN No. 1109

length ahead of the right wing tip and was continuously recorded by a single-cell manometer recorder. A continuous record was also obtained of the difference between the fuselage static-orifice pressure and the free-stream static pressure. Impact pressure of the free-air stream was obtained from the record of the shielded total- pressure tube and the free-stream static-pressure tube on the wing- tip boom. •-- •

The angle of attack of the thrust axis for the level flights was obtained by measuring the inclination of the thrust axis with an inclinometer. Cowl-flap angle was obtained with a calibrated electrical position indicator. The relation between cowl-flap angle and cowl-flap exit area is shown in figure 5.

Test and analysis procedure. - The analysis of the data was accomplished by comparing runs in which all of the conditions were maintained approximately constant except for the variables being investigated. The desired conditions could not always be maintained precisely constant but the pressures were generally little affected by the variations that occurred. Various combinations of flight variables were possible by lowering the landing flaps and by extending the landing gear, thus changing the drag of the airplane. A summary of flight conditions as well as computed propeller coefficients are given in table II; figure numbers for the curves showing the test results are also included. The thrust disk-loading coefficieiit of the propeller Tc was computed from brake horsepower, free-stream impact pressure qc, propeller-disk area S, and propeller efficiency T). Information from the Propeller Division of the Curtiss Wright Corporation was used to set up the propeller-performance curves (fig, 6) from which the propeller efficiency was determined.

In order to show the degree of stability during the flights, typical NACA pressure-cell records of free-stream impact pressure, the fuselage static-orifice pressure, and pressure altitude are shown in figure 7 for one flight run.

RESULTS AMD DISCUSSION

The discussion of the results is divided into four parts: (1) average recovery and circumferential distribution; (2) radial distribution; (3) longitudinal distribution; and (4) comparison of pressure-drop measurements. The engine coolingr-air pressures pre-_ sented herein are shown as a ratio of the measured pressure to free- stream impact pressure. This ratio for pressures in front of the engine will be referred to as "front recovery."

NACA Tlf No. 1109

Average Recovery and Circumferential Distribution

Effect of airplane speed. - The effect of changing airplane speed, as normally accomplished "by changing engine power, on average engine cooling-air pressures (average for nine cylinders of one row) for various stations in front of and behind each of the cylinder rows with cowl flaps closed is shown in figure 8. Increased airplane speed thus obtained is accompanied by changes in other flight variables that are dependent upon the airplane and propeller-performance characteristics. The front recovery, which was relatively low as compared to wind-tunnel tests of cowlings of the same general type, increased with airplane speed; an increase in airspeed from 185 to 255 miles per hour resulted in an increase in recovery from 0.67 to 0.77« The average rear pressures were affected by increased airplane speed approximately the same as were the front recoveries, therefore making the ratio Ap/q^. a constant at varied airplane speed for closed cowl flaps. This trend indicates that the cooling- air weight flow would increase only slightly for this installation with an increase in front recovery.

The effect of increased airplane speed on the circumferential pressure distribution at various locations in the nacelle with cowl flaps closed is presented in figure 9. The front pressures show an improvement in the pattern with an increase in airplane speed, which results from a larger increase in the pressures on the top of the engine than at the bottom. The improvement in distribution in front of the engine as well as the increase in average pressure recovery was the combined result of changes in thrust disk-loading coefficient and angle of attack of the thrust axis, which will be discussed later. The distribution downstream of the cylinders was not noticeably affected by the increase in speed for closed cowl flaps.

Effect of cowl-flap exit area. - The front recovery was affected only slightly by increasing the cowl-flap exit area at cruising conditions although it tended to decrease in front of the rear row (fig. 10). The rear pressures showed an average decrease of about 0.15 <!(,, with the pressures behind the front-row barrels being least affected. Consequently, the decrease in rear pressure with an increase in cowl-flap exit area greatly increased the Ap/q^. ratio across the engine. This result indicates that one way to increase in cooling-air weight flow is to decrease the flow losses at the rear of the engine and from the cowl exit without changing the cowl exit area.

The effect of cowl-flap exit area upon pressure distribution is shown in figure 11. The front-row baffle-entrance pressure distribution was affected very little by opening the flaps; whereas the rear-row baffle-entrance pressures were slightly decreased on the outboard side of the engine (cylinders 3 to 9). The pressures on

HACA Tit No. 1109

the inboard sido were possibly influenced "by the restriction formed "by the nacelle, wing, end fuselage. The distribution behind the engine was appreciably affected by opening the cowl flaps; the decrease in pressures was about 50 percent greater for the bottom cylinders in the region of the flaps than for the top cylinders. The low pressure behind the cylinders in the region of the flaps together with the relatively high front recovery for the bottom portion of the engine results in a larger pressure drop across these cylinders, even when the flaps are in the full-closed position (46 percent full-open exit area, fig. 5). This difference in cylinder cooling-air pressure drop would result in a temperature difference among cylinders for closed cowl flap, cruising operation of about 40° F, as calculated by the cooling-correlation equation established in reference 1; with the cowl flaps in full-open position a spread of approximately 50° F could be expected due to cooling-air flow distribution. From these calculations it is evident that the largest part of the temperature difference resulting from cooling-air flow distribution was caused by entrance conditions and the circumferential location of the cowl flaps, together with the fact that tnere was very little exit area from the cowling except through the flaps.

The effect of opening cowl flaps on cooling-air pressure recovery and distribution at a density altitude of 20,000 feet (figs. 12 and 13) was similar, in general, to that at a density altitude of 5000 feet (figs. 10 and 11) except that both the front and the rear pressures decreased slightly more at an altitude of 20,000 feet when the quantity of cooling-air flow through the engine was increased. The resulting Ap/q^, ratio for the various cowl- flap exit areas was, however, approximately the same for the two altitudes. -.

Effect of angle of attack of thrust axis. - The effect of increasing the angle of attack of the thrust axis upon pressure recovery and distribution is shown in figtu-es 14 and 15, respectively, for closed cowl flaps. All average pressures decreased because of air spillage over the top of the cowling. The increased spillage greatly decreased the front pressure available for cooling the top cylinders. Another contributing cause of the decreased pressures at the top of the cowling was the blanketing effect of the spinner at high angles of attack of the thrust axis. This decrease in the pressures in. front of the top cylinders may become more important at greater angles of attack such as are encountered in take-off, climb, or high-load conditions; in this event, the temperature distribution would be appreciably affected. The bottom pressures were less affected and in some cases were increased with increased angle of attack owing to improved entrance conditions at the bottom of the cowling. The rear pressure distribution remained essentially the same.

8

KACA TU No. 1109

The flow characteristics of the cooling air into the cowling vere verified by observing tufts attached in front of the engine at the entrance and on the inside of the cowling. The air flow was noted to be relatively steady in the bottom portion of the cowling and unsteady in the top portion where spillage was readily apparent. An adverse pressure gradient resulting from the abrupt break in the profile of the flow path at the rear of the spinnei- was indicated by tufts around the reduction-gear housing.

Effect of propeller thrust disk-loading coefficient. - The effect on cooling-air pressures of the thrust disk-loading coefficient Tc, which is indicative of the pressure rise across the propeller, is shown in figure 16. For open cowl flaps a decrease in front recovery of 0.10 resulted when the thrust disk-loading coefficient was increased 0.19. The increase in this coefficient was obtained by decreasing the impact pressure. It was concluded in reference 3, that the pressure available for cooling (cooling- air pressure drop, as used in reference 3) is a direct function of the thrust disk-loading coefficient. The increase in pressure drop with the corresponding increase in disk-loading coefficient (shown in reference 3) was largely a result of an increase in front recovery, especially for closed cowl flaps where a change of slip- stream velocity ha3 little effect on the cowl-exit pressure. The difference between the two sets of results is undoubtedly due to the differences in the respective installations. In the test installation used herein, the root section of the propeller with the cuffs appeared to be very ineffective and the nacelle-propeller diameter ratio was only 0.33. In other installations where the propeller-root section is more effective or the nacelle-propeller diameter ratio larger, an increase in thrust disk-loading coefficient would increase the pressure in front of the engine. The average rear pressures of this installation were decreased approximately the same as the front recovery, which resulted in relatively little change in the ratio Ap/q^ when the front recovery was changed although the cowl flaps were full open for these flights. _

The pressure-distribution pattern for the cylinder heads was only slightly affected by the changes in thrust disk-loading coefficient: whereas the top front-row cylinder-barrel pressures were decreased more than other barrel pressures at the low-speed high- thrust condition (fig. 17). This decrease indicates a large adverse pressure gradient on the top of the reduction-gear housing resulting in separation from the spinner. The pressure-distribution pattern in front of the rear-row barrels was unaffected by the poor flow characteristics on top of the engine.

NACA TN No. 1109

Effect of propeller speed. - A variation of the propeller speed and. consequently, cf the advance-diameter ratio v/ND had no substantial effect upon the pressure recovery and distribution at cruising power (figs. 13 and 19), although the recovery was increased slightly at propeller speeds of 100C to 1100 rpm. Apparently the variation in "blade angle had very little effect upon the pressure available for cooling in this installation and the rotation of air behind the propeller was not sufficiently changed to affect the distribution within the cowling. "

Comparison of baffle-entrance pressures on exhaust and intake sides of cylinders. - The effects of airplane speed, cowl-flap exit area, angle of attack of the thrust axis, and propeller speed on the difference between the baffle-entrance pressures on the intake and exhaust sides of the cylinder hoads and barrels are shown in figure 20. The baffle-entrance pressures on the intake and exhaust side of the front-row heads and barrels W6re very nearly equal in all cases, indicating no appreciable change with operating condi- t.lons. The rear-row head pressures, however, were low on the exhaust side and the barrel pressures were slightly high on the exhaust side. Cowl-flap exit area was the only variable that affected these pressure differences; the spread between the pressures of the two sides was increased as the cowl flaps were opened. This effect was less noticeable for the barrels where the pressure difference was small.

Kadial Distribution

The distribution patterns presented in the preceding section have Indicated that the radial distribution of total pressure in front of the engine and of static pressure at the rear of the engine vari.ed among cylinders; three front-rcw and three rear-row cylinders were selected to show this variation in the radial distribution at different locations. The locations chosen were the top of the engine, the cowl-flap region, and the bottom of the engine.

Representative plots of the radial pressure distribution at various airplane, speeds, cowl-flap exit areas, and angles of attack of the thrust axis are shown in figures 21, 22, and 25, respectively. Airplane speed and angle of attack had no appreciablu effect on the distribution either in front of or at the rtsar of the engine. Cowl- flap exit aroa had little effect upon the distribution upstream of the engine although It tended to become less uniform for the bottom cylinders as the cocling-air flow was increased because of. the flow characteristics of the air entering the bottom of the cowling.. Tho gradient of tho static pressures bt>hind the cylinders was increased with an increase in cowl-flap exit aroa, particularly for the_

10 _._.

NACA TN No. 1109

front-row cylinders. The difference in pressure gradients in front of and behind the individual cylinders, however, was greater than the change in gradient due to varied operating conditions.

The radial-distribution patterns indicate that distinct differences exist in the radial pressure distribution between the front- and rear-row cylinders. In this particular installation, the entrance pressures for the front-row cylinders were highest near the middle of the cylinder; whereas, for the rear-row cylinders the pressure was lowest near the middle with exception of the bottom cylinders where more stable flow into the cowling prevailed. The pressure distribution behind the engine cylinders was affected largely by the circumferential location of the cowl flaps as indicated by the large pressure gradient in the cowl-flap region.

When changes occur in the radial pressure distribution of a given installation or when like engines are placed in different nacelles that do not have the same radial distribution, the different engine air-flow conductivities and consequently the different mass-flow pressure-drop relations that will result are important considerations.

Longitudinal Distribution

The relations between the useful pressure drop across the cylinders and the entrance and exit losses of the cylinders have been indirectly shown by the patterns of circumferential and radial pressure distribution. These relations are, however, more conven- iently shown by plots of longitudinal distribution of pressure through the engine. Such curves are presented in figures 24 and 25 for closed and open cowl-flap positions, respectively. The distribution at three circumferential locations around the engine and the average distribution are included. The total pressures ahead of the cylinders are those measured by tubes at the baffle entrance. These tubes (on all front-row cylinders) were used for the pressures ahead of the engine and indicated pressures of approximately the same magnitudes as the shielded tubes in front of four cylinders of the engine. The use of these tubes prevented determination of the baffle-entrance losses to the front-row cylinders but these losses were undoubtedly small. The pressures directly behind the cylinders were measured by total-prossure tubes rather than static-pressure tubes in order that the exit losses might be evaluated. The rear- most pressures behind the engine were measured by static-pressure tubes behind the intake pipes where the velocity pressure was small; the differences between this pressure and the front-row baffle- entrance pressure is considered to be the total pressure drop~~äcross the engine installation.

11

NACA TN No. 1109

For closed ccrwl flaps the longitudinal distribution through the engine is similar at each of the three circumferential locations "because it is chiefly dependent 011 the absolute value of pressure drop across vany particular region of the engine (fig. 24). On the assumption that the "baffle-entrance losses are negligible, the average pressure drops across the front-row heads and barrels were 55 and 40 percent of the total pressure drop across the engine, respectively. The baffle-exit losses made up the remaining portion of the total. In the rear row, the entrance loss for both the heads and the barrels was about 20 percent of the total; the drops across the heads and barrels were 65 and 50 percent of the total, respectively, and the exit losses were 15 and 30 percent, respectively. From a comparison of the two rows, it is noted that, regardless of the rear-row entrance losses, the useful pressure drop across the rear row is roughly 20 percent greater than that across the front row. This difference in useful pressure drop across the two rows, if it is assumed to be a reliable indication of distribution of cooling-air weight flow, would result in the front-row cylinders running 10° to 30° F hotter than the rear-row cylinders, dependent upon operating conditions. For an engine that develops a greater amount of power in the front row than in the rear row, as reported by the Army Air Forces in 1943 (Memo. Rep. Ser. No. 57-503-858), this cooling-air flow distribution between rows could be of considerable detriment to efficient operation. - --— •--•

The longitudinal distribution for open cowl flaps is shown in figure 25. The distribution between the front and rear row is very nearly the same as for closed flaps (fig. 24); the difference between the head and barrel pressure drops, however, became slightly larger when the flaps were opened.

Comparison of Pressure-Drop Measurements

A large number of different types of pressure tube at various locations hr.ve been used in test-stand, wind-tunnel, and flight tests as an index of cooling-air flow through an air-cooled engine. The location of many of the pressure tubes was duplicated in the present tests enabling a comparison of pressure-drop measureinents. This comparison may be used to facilitate correlation of various cooling investigations that have employed different methods of measuring pressure drop. The results are tabulated in table III by listing the various pressure-drop recoveries for open and closed cowl flaps and by showing the .relation between the different pressure drops by comparing them with the pressure drop used in reference 1 (Ap^), The table includes two general types of cooling-air pressure-drop measurement. Methods 1 to 5 show the difference

12

WACA TN Hb. 1109

"between average entrance pressure of front-row cylinders and average exit pressure of rear-row cylinders; this type of measurement includes the losses in the entrance passages to the rear-row cylinders and in the exit passages from the front-row cylinders. Methods 6 to 9 show the difference between the average entrance and e:cit pressures of the individual cylinders, thereby excluding the entrance- and exit-passage losses. As shown in table III, a large difference exists among the various pressure-drop measurements. Across the heads, the largest indicated pressure drop is almost twice as great as the smallest one; whereas across the barrels, the difference is larger. The relation between the various pressure drops for this installation was little affected by the cowl-flap position, although, opening the cowl flaps increased" the value of the Ap/q^ ratio roughly 60 percent. The large differences in the values of pressure drop obtained by different methods of measurement and the effect of different installations on engine cooling-air distribution indicates that good coin-elation of the cooling results of like engines in different installations cannot be expected unless the instrumentation and installation differences are taken into account.

Because of the difficulty in accurately measuring ceoling-air weight flow in a flight investigation, a qualitative comparison of the reliability of the various pressure-drop methods was impossible, even though, large differences emong the various measurements jrore shown. The pre3sure-drop method used in reference 1 (Ap-,) gave the best total engine cooling correlation; however, this comparison is dependent on the accuracy of the correlation procedure in accounting for differences in cooling variables other than cooling- air weight flow. Consequently, this procedure is not considered sufficiently conclusive for malcing a qualitative comparison of pressure-drop measurements.

SUMMARY OF RESULTS

From the flight investigation of tlie engine cooling-air pressure recovery and distribution of a two-row radial engine enclosed in a low-inlet-velocity cowling, the following results were obtained;

1. The average front pressure recovery, which was relatively low for this engine installation, increased with an increase in airplane speed during normal level flight. —

2. The pressure drops across the front-row cylinder heads and barrels for closed cowl flaps were 55 and 40 percent, respectively,

13

HACA TH No. 1109

of the difference "between the pressures front and rear of the engine; the pressure drops across the rear-row heads and "barrels were 65 and 50 percent, respectively.

3. The static pressures "behind the heads were lower than those "behind the "barrels; this difference increased when the "cowl flaps were opened, especially for the front-row cylinders in the cowl-flap region. This change in radial distribution with operating conditions was smaller than the difference "between individual cylinders.

4. The pressure distribution in front of the engine had little effect upon the distribution behind the engine; however, an increase in average front pressure recovery resulted in almost as large an increase in average rear pressure.

5. A change in pressure at the rear of the engine accomplished by varying the cowl-flap area had little effect on the pressure recovery and distribution in front of the engine.

6. The general pattern of the circumferential distribution behind the engine was chiefly determined by the circumferential" location of the cowl flaps.

7. An increase in angle of attack of the thrust axis decreased the front recovery at the top of the engine because of spillage from the cowling and separation from the spinner; the air flow into the bottom of the cowling remained relatively steady. —

8. An increase in propeller thrust disk-loading coefficient decreased the average front recovery for this installation.

9. The speed of the propeller bad little effect upon the pressure recovery and no effect on the distribution at cruising power.

10. Large differences, sometimes amounting to 100 percent, were obtained among the results indicated by various methods of measuring pressure drop across the engine.

C0HCLUDIE& HBMAHKB —

The results of these tests indicate that an important consideration in the design of cowlings and cowl flaps should be the obtaining of good distribution of cooling air as well as minimum drag for the installation; the results further show that the cooling-air flow" distribution and, consequently, the temperature-limited performance of a given engine installation is considerably affected by cowl- entrance conditions and circumferential location of the cowl flaps.

14

NACA TN No. 1109

The fact that these teats showed that the front recovery- decreased -with an. increase in propeller thrust disk-loading coefficient proTidos additional evidence that the recovery is greatly- affected by the combined propeller-nacelle design. Also of significance is that a large increase in front recovery resulted in a similar increase in rear pressure, indicating that an increase in the front recovery of an air-cooled engine is not always an effective method of increasing the cooling-air weight flow.

Aircraft Engine Research Laboratory, National Advisory Committee for Aeronautics,

Cleveland, Ohio, December 3, 1945.

REFERENCES

1. Bell, E. Barton, Morgan, James E., Disher, John H., and Mercer, Jack R.: Flight Investigation of the Cooling Characteristics of a Two-Row Radial Engine Installation, I - Cooling Correla- tion. NACA TN No. 1092, 1946.

2. Pinkel, Benjamin: Heat-Transfer Processes in Air-Cooled Engine Cylinders. NACA Rep. No. 612, 1S3B. -:-.

5. Stickle, George ¥., Naiman, Irven, and Crigler, John L.: Pres- sure Available for Cooling with Cowling Flaps. NACA Rep. No. 720, 1941.

15

TABLE I - ES0I1IE COOLIHO-AIH PRESSURE-TUBE ISSTALLATIOH > O > Free

aura tube

t»l Relativ« location Type

Circumferential location

Cylinder« Position relative to aylInder

Axial loaatlon B&dlal location fKOD ajllnder- flanga baa«

(in.)

T

4 4

«t

4

4 «*

Hhl

"hi

%8i

"net

"bei

•bS.

•M

\i

=b4

Pb4

»h6

PM

*h7

Pb7

"ha

Front of haad on rake

Front of barrel on rake

Between fin and baffle at head-baffle antranee

-do-

Shielded total head

Total bead

-do

2,6,31,16

-do-

Center of oyllnder

-do-

All

-do--

«-de—- -do-

Between fin and baffle at barrel-baffle entranoe

—do—

-do-

-do-

Front of oowl sealing ring

Rear of head on rain

-do— ---------—.

Rear of barrel on rake

.do———

Behind oowl sealing ring

Rear of barrel between flange and fin*

Downatraaa of engine (heads)

Downstrea» of engine (barrels)

Downttwe* of angina

Baffle tap

Total head

Clo*«d-end atatla

Total head

Cloaed-end atatie

Open-end atatie

-do

1,4,5,7.6, 10,11,14, 15,17,18

All

1,4,8,7.8, 10,11,14, 15,17,18

1,7,18

All

Intake aid«

Canter of oyllnder

Exhaust aid«

Intake aide

Exhauat aide

Center of oylInder

—do-

5g upstream of front-row baffle entrance

•do -——— —.

5/lfl downstream of baffle-cntronoe owl

-do-

•do-

-do-

-do-

-d<r- -do-

On baffle butting against sealing ring

7/8 downstream of head fine

-do — -

•do—

-do-

—do—

-do-

7/8 downetream of barrel fine

-do-

-do—

-do—

—-do——

-«0

-do-

- do-

Hear row

-do-

In baffle-exit curl, Intake side

—do

Closed-end atatla 8,11,17 Behind aharge-alr Intake pipe

1/3 behind head sealing barrio

1/16 behind oyllnder barrel

At baffle exit

•do--—----——---—— •

E behind Intake pipe .

O

'See figure 3 for explanation of synbols and aubaorlpts. NATIONAL ADVISORY

COMMITTEE FOR AERONAUTICS o>

TABLE II - SUMMARY OF PLIGHT COHDITIOHS AHD COMPUTED PROPELLER COEFFICIENTS

Airplane conditions Engine conditions Propeller coefficients

Density Pressure Impact Angle of Cowl-flap Brake Engine Speed- Advance- Thrust disk- of free altitude pressure attack of angle

(deg) horse- speed power diameter loading

Figure air (ft) (in. water) thrust power (rpu) coeffi- ratio coefficient (slug/ axis cient V/ND Tc cu ft) (de«) ca

8,0» 0.00205 3455 16.6 4.5 Closed 783 2400 1.876 1.066 0.102 20,21 .00206 5494 22.0 3.3 —do 1012 2392 2.057 1.230 .087

.00205 5710 28.9 1.7 —do 1267 2400 2.210 1.3B0 .072 «1 1 1 C .*** _. ww w 1 K4Q •fabTa ft 041« O.Srtl 1.489 «072

10,11, 0.00206 3178 22.8 2.7 Closed 1008 2408 ' 2.083 1.240 0.082 20,22 .00206 3196 21.8 2.6 . 14.5 1017 2414 2.036 1.210 .069

.00206 3186 20.6 3.7 27.0 1019 2420 1.981 1.175 ,097

.00206 3208 19.7 3.6 Open 1024 2420 1.934 1.148 .104 12,IS 0.00129 18257 14.S 5.1 Closed 825 2420 2.001 1.255 0.101

niM O0 1 AVAO 1 C 4 .4 e ten oi a OiAA © ft«;© 1 %Of\ 1 ITA n.w ^w. w 0<VU Arf~WI W b . WM *. «VMV

.00127 18289 14.3 5.0 30.8 916 2400 1.959 1.261 .119

.00127 16240 14.1 5.1 Open 916 2400 1.950 1.250 .121 fit«« «u

A a WJLWi»VU mil OArtrt o.njw 1 .OX ft n.nfls

20,23 .00206 3494 22.0 3.3 —do—- 1012 2392 2.057 1.230 .087 .00206 4820 18.6 5.1 do-- — 804 2414 1.981 1.123 .088

16,1? 0,00207 5050 10.2 Open 1UCÜ 1 AAC ft £A1 A OrrA Vet- *"•

.00206 5112 14.0 1.8 —do 1023 2406 1.621 .970 .175

.00206 5133 17.0 1.8 do---» 1024 2408 1.792 1.072 .131

.00206 5176 22.4 3.1 ---do--»« 1023 2404 2.055 1.232 .086 16,19, 0.00206 3866 21.7 3.4 Closed 1013 1806 2.273 1.610 0.089 20 .00205 3886 21.5 3.4 —do—— 1010 1996 2.117 1.457 .087

VOQfi Ol Q 11 .A «*«« WW~ — WBS lflTJ oonn O-Qru 1 -ÄA7 .088 .00205 3886 22.3 3.1 —do—— 1020 2398 2.063 1.234 - .069 .00205 3886 20.8 3,4 — dO- 1009 2584 1.933 1.105 .087

a* 9^ MA OLIO «CO

rn _ . I UJ.UJJOU

1 jtnA nj no 61UD CfUDV 1 OJ rt n noo

W. WWJw

26 0.00206 3208 19.7 3.6 Open 1024 2420 1.934 1.14B 0.104

>

nniivnib

COMMITTEE FOR AERONAUTICS

O

N AC A TN NO. 110 9 18 TABLE III - COMPARISON OP VARIOUS PRESSURE-DROP MEASUREMENTS

[Altitude, 5000 ft]

Ap

Pressure- drop method

4p/q0 ap/APi

Cowl flaps closed

Cowl flaps full open

Cowl flaps closed

Cowl flaps full open

Head

»1 fHh3i J *la2t\ ( 0.28 0.45 1.00 1.00 V s )tr

m ^^„ 2 fHh2i + Hh2t\ , (p » 0.33 o.so 1.18 1.11 V 2 ;fp

(IWrr

3 fHh21 + Hh2t\ ' ,r . 0.29 0.45 1.04 1.00 V 2 A. (h/,rr

4 f Hh2I + «h2t\ f^& + ^ 0.31 0.48 1.11 1.07 V 2 >,r " V 2 yrr

S fHh2i + HhBt^ f» , 0.25 0.42 0.89 0.93 V 2 >>fr ^4)^

6 (^h21 + Hh2tN ,p . 0.22 0.34 0.79 0.76 V 2 ;ae (P^Jae

7 fHh2i + Hh2t\ ,_ . 0.30 0.46 1.07 1.02 V 2 ;ae (phb>ae

e fV1! Hh?y\ - (RJI 0.18 0.30 0.64 0.67 V 2 Vae ( h4)ae

9 bHhl ' P^a. 0.24 0.37 0.86 0.82

Barrel

ftl Hb21fr " Pt*rr 0.24 0.40 1.00 1.00

2 Hb2lfr - Pb6pr 0.26 0.42 1.08 1.05

3 W ' Pb7rr 0.27 0.43 1.12 1.07

4 W -^^i 0.26 0.42 1.06 1.05

5 Hb2ifr - Hb4rp . 0.17 0.31 0.71 0.77

6 ^"ae " ^*ae

0.16 0.27 0.67 0.68

7 Hb2iae " Pb6ae 0.16 0.26 0.67 0.65

8 Hb2Iae " HMM 0.10 0.19 0.42 0.48

9 BHbl " Pb4ae 0.19 0.30 0.79 0.75

aPressure-drop method used in reference 1. °Front of cylinders 2, 6, 11, and 16 only.


N ACA TN No. I »09 Fig. [

Figure I. - Test-engine installation,

28-ftl

Chirge-air ducts

Front viev

fiflore 2. -

Cuff

a*.*y >;"

It« tMf\

Spinn

Cow» ffap»

Side view

Sketch Of low.Inl.t i . „ N*T'0«AL ADVISORY low-fnlet-velocity coding and propeller «i • F°* Ä"•s

peMer »P'nner and cuffs.

a: o

Q

• 1

03

N)

Seallng-rlag baffle

*»T t\d*

Symbols: total pressure

criptt: head barrel

Tl

CO

P

I to 3 longitudinal station relative to cylinder

i Intake 9lde of cylinder e »haust side of cylinder t top of cylinder

> >

h2a


(a) Front pressure tubes..

Figure 3. - Front- and rear-row cylinders showing pressure-tube installation. o

et 1 J

*$>J I •

z > O >

o

o

Symbols: H total pretnre t static presaure

Subscr ipta: h head ' t> barrel

4 to 8 longitudinal station relative to cylinder


F i gu re 3. t I on .

- Cone Iuded.

rb7 iiw" ' on -'b6

(bl Rear pressure tubes.

Front- and raar-row cylinders showlnjg pressure tube installa-

[ I I I

NACA single-cell recording «anometer (f ree-streaai static pressure - fuselage static pressure;

Wing boon -j NACA single-cell recording •anoneter (impact pressore}

^Q7 \

Total-pressure tube-*

100-tube sei ector val ve

100-tube I i quid •tanoaeter

T1

to

> >


Figure •*. - Diagrammatic sketch of system for measuring pressures. o

i.ti NATIONAL ADVISORY

COMMITTEE FOR AERONAUTICS

y 2.6

2.4 • Cow )-fl »P t Kit arei t » 1 ft

^

<*-

<r 2.2 •

*

a /

A

M a

• -^C - 1.8 •

^

\ -Cow l-fl ip exit are« » * o o

/

per :ent of full -op«n area

1.6 j /

,S y Full-closed

Fill-open,/ / * exit area —1—

/

/

1.4

f J /

1.2 /

0 4 8 12 16 20 24 28 Co*I-flap angle, deg

Figure 5. - Relation between cowl-flap angle and exit area.

32

«

c *> e. o •

100 z

90

80

70

60

50

c

L. a

a. a

o o

36 40

o >

o

CO

Ul

Fig N ACA TN NO. I 109

O)

Ü

0) o o

<v s o I

•o a>

4.2

3.8

3.4

3.0

2.6

2.2

1.8

1.4

.8 • 1.2 Advance-diameter

2.0 ratio, V/ND

Figure 6. - Computed performance installation. Propellers: blaue » c ^ t. i u n , i. i a r * diameter, I 3± feet; number of blades, 4; nacelle- propeller diameter ratio, 0.33.

__ curves for propeller blade section, Clark Y;

N AC A TN No. I 109 Fi g

(A « a W k & 01

u *

*> u

22

21

20

19

« ii ii ii

a a m ' — as o «I IA • «

3400

- 3200

3000


l/v\

/\ / -^> /> s~+ A A r \ t

^ J V Av

^/ ' \ A -^

,«s y \ / V

„-=1

20 40 SO 80 I0O Time, tec

120 140 160 180

Figure 7. - Typical record of free-stream impact pressure, fuselage stat Ic-orlfIce pressure, and pressure altitude.

Head pressures

°l»h2i+«li2t>/2 oPh5 A".4

•*rn Air flow

Barrel pressures

• "oil r Hb4

* Pb4

o Pb7

X'bB

•60 220 260 300 ISO 320 260 300 Indicated airspeed, «ph

Figures. - Effect of airplane speed on average «ngine cooling-air pressures during no rail level flight. Density altitude, 5000 feat; cowl flaps, closed; propeller speed, 1200 rpm.

Tl

to

00

> o >'

»I

N ACA TN No. I 109 Fig. 9a

2 > 9 a.

o

to

.6

.3


O •

o

Indicated airspeed (nph)

182 209 240 25«

2 4 6 8 10 12 14 16 Cylinder, front row

(a) Cyl Indsr-head pressures.

i^j^S

18 3 6 7, 0 II 13 IS 17 Cylinder, rear row

Figure 0. - Effect of airplane speed on circumferential pressure distribution during normal level flight. Density altitude, 5000 feet; cowl flaps, closed; propeller speed, 1200 rpn.

Fig. 9b N ACA TN No. I I09

*


.ü

Ed- ]£-" . / *-"i I • 6 • Ji

"* pb6 i Air flö»

• 0 i >*---1 > < s >^ > . / ! ^ ] 1 ^

* \ « i

[ Nu !==( i*1

^ M

^

i.

.ti "Vj l-J 5-^« 1 i

• C

• 7 * t=x i < > • 1—! j—i r^ t

I )—<; \ t- V —H • 6 \ **** } s ) 's M J-* i ^t *H ]—i la i w *3 ] -b c ^ ^ ^=5 r^ N r

• 7

1 ^£ >_ .

FS ^ SN , J v-^ >-< r f Si n R i fcj >~^ 4 f> J w> Ö ? N *Ü I—i !—5

j^ i

\ l >

' -< »--=

, 7

*~< fT-i kS < *=M r

\ F=* *=< 3=^ ^ s 1 < I—c >

• 6

j—t r3 P=« r •

Indicated ai rspeed (mph)

O 182 D 209 A 240 O 256

• 6

• u \ *=* »^ rd t=3 I

^ !=* tJ &

2 4 6 8 10 12 14 16 IB Cylinder, front row

(b) Cylinder-barrel pressures.

I 3 S 7 9 II 13 16 17 Cy I Inder, rear row

igure 9. - Concluded. Effect of airplane speed on circumferential pressure distribution during normal level flight. Density altitude, 5000 feet; cowl flaps closed; propeller speed, 1200 rpm.

e

•

o> o u

c

Z > n >

60 80 100 40 80 80 Cowl-flap exit area, percent of full open

Figur« 10. - Effect of cowl-flap exit area on average engine coollng-alr pressures. Density altitude, 5000 feet; free-stream Impact pressure, 20 to 23 Inches water; angle of attack of thrust axis, 2.6° to 3.7°; thrust dlsk-toadlng coefficient 0.08 to 0.10; propeller »peed, 1200 rpn.

T|

Fig. I la N ACA TN No. | I 09

? o c u

1?

NATIONAL ADV 1SORY . COMMITTEE FOR AERONAUTICS Due w

A SL

7 I 1 n ^ £ Ä N i - 1 \ =J r i > 0

• W Air flo

8

7 1 »H pN ;> ^fefe* r^ IS & s I 6 y ^ r

5

y. <

1 **

>—< £r >-^

&

1 J * t fc

i—< ^ 1——; L

£•*. k ^

1. -.,

i r^ s. .^< 1 1

kS A t J ^ i—( >

6$ s ^""c

... • >s^

^

1—I .—;

J

"*< *^ ^ s

# ^ >'

\ > ^ V 1 * >

A \ 0 < »» /

X < L, ^

E *^c c r' i * ft ! N 1 • -^

1 3 <

^ k i—\ ] <

^ ^ \ I—t 2 *

N >—< t yf s S

.—* ? P* \ r

8

4

3

Cowl-flap exit area (percent of fu11 open)

O 40 D 71 A 86 O 100

i k^ • ( ^

i i

r \ & ^ 'A < y-*l ***-< «^

6 7 S II 13 16 Cylinder, rear row

17 0 2 4 6 S 10 12 14 16 18 Cylinder, front row

(a) Cylinder-head preseurea. Figure II. - Effect of covl-flap exit area on circumferential pressure distribution. Density

altitude, 5000 feet; free-stream impact pressure, 20 to 23 inches water; angle of attack of thrust axis, 2.6° to 3.7°; thrust disk-loading coefficient, 0.06 to 0.10; propeller speed, 1200 rpm.

N ACA TN No . 1109 Fig. lib

e er = > 8 <e

0 2 4 6 8 10 IS 14 16 18 I 3 S 7 9 11 13 15 17 Cylinder, front row Cylinder, rear ro»

(b) Cylindtr-barrel pressures.

Figure II. - Concluded. Effect of cowl-flap exit area on circumferential pressure distribution. Density altitude, 5000 feet; free-stream Impact pressure; 20 to 23 inches water; angle of attack of thrust axis, 2.6° to 3.7°; thrust disk-loading coefficient, 0.08 to 0.10; propeller speed, I 20 0 rpm.

a o u

<a >

40 60 80 100 40 SO 80 Cowl-flap exit area, percent of full open

100

Head preaaire*

AHh4

-=

'—~~-~i -*"

i

3 _,^^--__

—V

9 1

I

1 l<= I" ~1

A r flow

Btrra I pressures

v Hb2j

r Hb4 4 Di j

> Pb7

*»»b6 > >

Figure 12. - Effect of cowl-flap exit area on average engine coollng-alr pressures. Density altitude, 20,000 feet; free~«tream Impact pressure, 14 to 16 Inches water; angle of attack of thrust axis, 4.5° to 5.1°; thrust disk-loading coefficient, 0.10 to 0.12; propeller speed, 1200 rpn.

o

NACA TN No. 1109 Fig. 13a

^ ^~? .5

.4

.3

.2

. I

i>—<> \z ]§£=>>

.4

.3 u

> .2

. I

.4

.3

.2

. I

Co«I-flap exit area (percent of fuI I open)

O 46

O 72 A 88 o loo

i^

-^.

^ [i—ri—1> vT

i^!

id^^Z. A**\


2 4 0 8 10 12 14 16 18 Cy I Inder, front row

(a) Cylinder-head pressures.

3 5 7 9 II 13 16 17 Cylinder, rear row

Figure 13. - Effect of cowl-flap exit area on circumferential pressure distribution. Density altitude, 20.000 feet; free-stream impact pressure, 14 to 16 inches water; angle of attack of thrust axis, 4.S0 to 5.1°; thrust disk-loading coefficient, 0.10 to 0,12; propeller speed 1200 rpa.

Fig. I 3 b NACA TN No. I I 09

^ > e a 8

*

^

Cowl-flap exit area (percent of ful I open)

O 46 a 72 A 86 O 100

<§^SS^^;

Air flow

7 5==!^

^sr ^:


3 0 7 9 II 13 Cylinder, rear row

8 4 6 8 10 12 14 16 18 I 3 0 7 9 11 13 IS 17 Cylinder, front row

(b) Cylinder-barrel pressures. Figure 13. - Concluded. Effect of cowl-flap exit area on circumferential

pressure distribution. Density altitude, 20,000 feet; free-stream Ira- pact pressure, 14 to 16 inches water; angle of attack of thrust axis, 4.5° to 3.1°; thrust disk-loading coefficient, 0.10 to 0.12; propeller speed I 200 rpm.

T «i

«a i at c

o

« a a

>

> O >

o

o

4 6 0 2 4 Angle of attack of thrnt axis, deg

. - Effect of angle of attack of thrust axis on average angln« cooling-air pressures, altitude, 6000 ft«t; free-atraan lapact pressure, 19 to 22 Inches water; cowl flap«.

Figur« 14. Density i , , _ . closed; thrust disk-loading coefficient, 0.09; propeller spaed, 1200 rpa,

Fig . I5a HACA TN No. I 109

.8 u or "i- .7 x

.6

^ .7

5 .8

*" .6

^ -6 >

.3

.0

.5

.4

.3

ii^L -^S§

N

<P*'

Angle of attack of thrust axis

(da0) O 0.9 D 3.3 A 6.1

^

ii- i|bJ=fc^fefe|^S?

2 4 6 8 10 12 14 10 18 Cylinder, front row

(a) Cyl inder-head pressure».

Air flow

j^bb£=L£W

—ik


3 S 7 8 II 13 15 17 Cylinder, rear row

Figure 16. - Effect of angle of attack of thrust axis on circumferential pressure distribution. Density altitude, 5000 feet; free-stream impact pressure, 19 to 22 inches water; cowl flaps, closed; thrust disk-loading coefficient, 0.09; propeller speed, 1200 rpm.

NAC A TN No, 09 Fig. I 5b

= >

«f

(b)

Figure 15.


1^3 !!

Angle of attack of < thrust axis

(dag) O 0.0 a 3.3 A S.I

2 4 6 8 10 12 14 Cylinder, front row

Cylindar-barral pressures.

16 18 3 B 7 9 11 13 Cylindsr, rear row

IS 17

Concluded. Effect of angle of attack of thrust axis on circumferential pressure distribution. Density altitude, 5000 feet; free-stream impact pressure, 19 to 22 inches water; cowl flaps, closed; thrust disk-loadtng coefficient, 0.09; propeller speed, 1200 rpm.

9

o

c ai c

« >

.20 .30 0 .10 .20 Thrust ditk»loading coefficient, Tc

Figure 16. - Effect of thrust disk-load!ng coefficient on auftrage engine cooling-air pressures. Density altitude, 9000 feet; free-stream Impact pressure, 10 to 22 inches water; cowl flaps, open; angle of attack of thrust axis, 1.8° to 3.7°; propeller speed, 1200 rpe.

> >

o

o

N AC A TN No. I I 09 Fig. !7a

.8

. 7

.6

.5

1

.7

.6

.9

.4

1

.5

.4

.3

.2

. 1

1

.0

.4

.3

.2

. 1

1

.3

.2

. 1

0

1

.3

.2

. i

0


u ^= * >», "h2t . Hk 11 .

-7fe^Z' -"1>5

-"1,7 o

E^- (T> "^s

^ i

j f ^Ph4

I

z

«

u Air floii

CW [

^ !?• N P2 £=i 1 s v.. JJ s <fi s 3 b> CM

Z

> > 3 t N ^ >-M ^

^ !

Sj / N < >

z

i fej &M *-«i ^rf 1 >- r^ & !

o < >—< S >—i

>—< S bi V t d i—i i J pq ! > * /- V ^ < 3 U^ S pa

iM-r

£-< > V

« a—( l f )

o >—5 ^

i < ^ „ <f i V c

a. 1 ;i ^ s i A F i

f X 1—< > i fcs ¥ ^ i

i fc* i ] \ hs i

^ I J >

i ^

^ \ /

i

/ $

Ö \ d V / < N <? b i t x > CL N > J £3 / N S >



r- * >< > \ i=*

1 <• s < ^ '

a. < >^

2 4 6 8 10 12 14 16 18 Cy I Inder, front row

(a) Cylinder-head pressures.

3 6 7 9 II 13 Cy• Inder, rear row

16 17

Figure 17. - Effect of thrust disk-loading.coefficient on ci rcunferen tla.1 pressure distribution. Density altitude, 5000 feet; free-stream impact pressure, 10 to 22 inches water; cowl flaps, open; angle of attack of thrust axis, 1.6° to 3.7°; propeller speed, 1200 rpm.

Fig. 17 b N ACA TN NO. I 109

.2 ef

2

NATIONAL ADVISORY 1

.8

!*• . 7

i fe* ==j F^ i»s ^ =f 1 "^ r y N »

** pb6

• •^ h-i ifc-f 1—i Air flow

—i tss i 1

/ 's 1—J •—* 1 ^ ^ H s^. • 6

< S >

< f \ > <| * ^

fcs t=s ^5 1

. 5 s >

. 4

1

.6 ( t 1 i >L , i

f~^ t=5i y - • 6 < >—< >

t < ^

i-—«

—' i—< > t ^ A T^ 1 .4 ,

> ^ s A M N H P-2

. 3 *s y^ \ • •^

.5

.4

.3

i \ <

. & b* ^ e5. . < ^ \ ö 3 P £-< >

l >=< *—? V

6

* & 1 !» , Iw ,i H H > % Q

A *—wf

""—J V f, >. / <. r-*= *—•< 7 s V r— •*< ''-r > ( N ^ 1

2 < *-< H *=? H l=* O >

Thrust disk-loading coefficient, Te

O 0.086 • . 131 A . 175 O .274

••i i '

)—< >--< >- 7 c ) 1 i i 3 1 5 f 7 Cylinder, front rov

(b) Cylinder-barrel pressures. Cylinder, rear row

Figure 17. - Concluded. Effect of thrust disk-loading coefficient on circumferential pressure distribution. Density altitude, 5000 feet; free-stream impact pressure, 10 to 22 inches water; cowl flaps, open; angle of attack of thrust axis, 1.8° to 3.7°; propeller speed, 120 0 rpm. • - —•

u 9

% Ul

O o o ID c

m

V >

900 MOO 1300 900 1100 1300 propeller speed (rpm)

Figure 18. - Effect of propeller speed on average engine cooling-air pressures. Density altitude, 5000 feet;, free-streaa (»pact pressure, 21 to 22 Inches «rntar; cowl flaps, closed; »ngla of attack of ttirust axis, 3.1° to 5.4°; thrust disk-loading coefficient, 0.09.

>

>

o

-n CO

00

Fig. I 9a N ACA TN NO. I 109

CM

CM

8 „ u cr > >

NATIONAL ADVISORY cs tUMMi i i ct rvt% Htnun

Hh9l . ZT

Ph5 -Pn7

8

7 Hhl-/|- j|^~Hh4

1 I Ph4

0

1

8 Ur flow

.. ^ B^ H >•!-{ I, 7

e

s

4

i N M H r feaj M ) < S »*H ^ I ^

7

e

e

4 i

^ 1 1 M »

6

i ? ^ ) 6

4

3

/ \ < k. J M > i 1 1 ^ r

5

.4

3

2

i V—Ti i u \ j N 1 ^ f \ s*

> > ¥* F"^ 1 s

^ I ^

\ r Propeller speed

(rppo) O 903 Q 998 A 1100 •> 1199 V 1292

S

6

h» 3 r Hß M I 3

•* «^ N M fed M W*

1 1 i I 4 e i 9 1 0 1 2 1 4 1 6 ( 8 1 3 & f i I i i 3 1 8 i 7

Cyl Inder, front row (a) Cylinder-head pressures.

Cylinder, rear row

Figure 19. - Effect of propeller speed on circumferential pressure distribution. Density altitude, 5000 feet; free-stream Impact pressure, 21 to 22 inches water; cowl flaps, closed; angle of attack of thrust axis, 3.1° to 3.4°; thrust disk-loading coefficient, 0.09.

N AC A TN No. I 109 Fig . I9b

1

(b)


• 8

.7 1 § s^fn HM2' "\1 t^-^ • 6

.6 •l'-jr ""* »b4

aa »bS Air flow

.7

.6

.6

,d to is fc* 1 r I *=5 I i M > r

.7

.6

.6

.4

.3

< *

1

.7

.6

• a

.4

.3

'

- IN H .7

.6

.5

.4

.3

.2

j L s fc=i N gs r* N

1 ^ r^

Propeller speed (rpm)

O 903 a 998 A 1100 O 1199 V 1292

.5

.4

.3

<N ei ^ I «^ M keä

• 2

8 10 12 14 16 18 Cylinder, front rov

Cylinder-barrel pressures.

3 S 7 9 II 13 Cylinder, rear row

IS 17

Figure 19. - Concluded. Effect of propeller speed on circumferential pressure distribution. Density altitude, 5000 feet; free-stream impact pressure, 21 to 22 inches water; cowl flaps, closed; angle of attack of thrust axis, 3.1° to 3.4°; thrust disk-loading coefficient, 0.09.

|n»act ereusre. In. water K-6 - 33

Ifcntst disk-loading coefficient Q.07 - 0.10

Angle of attack of tferast axis, deg 1.5 - 4.5

Coal-flap positioa Closed

Propeller «peed, ran 1200

20-23

o.oa - 0.10

2^-3.7

1200

19-22

D.09

Closed

121»

21-22

0.00

3.1 - 3.4

to

••• •

Figure aad


20

10

0 _I1— •o— » » 1 L

0

( i) rroa t-ra » «* Mad tr 1 oadi ,

to

0

10

,^— 1 :—*

( >> :roa L-rc • cy Mid tr t irrt Is.

JO

20 o—

10 ** F— -

i :j ioar -roi cyi ind< r k< ad».

10

0

10 ( i) Mar -ro« ejl ladt r ba rral ••

•

ISO 200 240 260 Indicated airspeed, epn

> n >

100 AegTe of attack of thrust ails, des

900 1100 1300 Propeller speed, rpa

40 60 80 Coal-flap eilt area, percent of full open

20. - Effect of flieht variables on the difference between baffle-entrance pretty res of Intake aad exhaust aides of cylinder heads barrels. Density altltede, 9000 feet. O

«

•a c

a id

u

m at

Fl

.2 .8 .6 .4 Engine cooling-air pranurt, H or p

qe gure 21. - Effect of airplane speed on radial pressure distribution during normal level flight. Density altitude, 5000 feet; cowl flaps, closed; propeller, speed, 1200 rpm.

> o >

o

o

Tl

CO

IO

16

12

O a

4

0

16

12

8

4

0

16

12

8

4

0

FIgure feet; 3.7°;

4«

C

ID TO

•*- o «I «I «

-0

e o

•T3


C R> ^ />

:, i /

'ft

/

i i

I 1 A 3,f *. /I

Cy) inder 2

-Jfifl i 1 F * <

f i / i

i

/

•y J ;> -T^^

^ '

//

Cylinder 6

\ k %r*o

**h ' (

rrV

Cylinder 10

•&>

M t vk

PI? n ~T7

*

Cy1inder I

<fa cs

Hi

CyI inder 7

# >J.v. v

Cylinder 11

—r

u ;(S

Cylinder location [Front view)

2 '

10 II

Air flow

Front of cylinder Rear of cylinder

Cowl-flap exit area (percent of full open}

o 46 (closed) a 71 a 85 O !Q0 (fa!! open)

.4 .2 .8 .6 Engine cooling-air pressure.

.4 H or p

22, - Effect of cowl-flap exit are» on radial pressure distribution. Density altitude, 5000 free-strea» lapect prrtsure, 20 to 23 Inch« water; angle of attack of thrust axis, 2.6° to thrust disk-load in« coefficient, 0.08 to 0.10; propeller speed, 1200 rpi«.

ro

> >

o

a at

«

ff a

m •a •

16

i n


0 J & .V /

1 0

A H 1

4

0 i • i k

Cylinder 2 ' 10

P P i p? f 12

/ // H I

1 r .' 8

,,< /

4

0

16

12

<(

* ?

•*

Cylinder 8

pf1

f 8 \ m

V 4 • ^ r 'A

0 Cyllndtr 10

.8 .ft .4 .2 Engine cooling-air pressure, H °r *

Ac

Figure 23. - Effect of angle of attack of thrust axis on radial pressure distribution. Density altitude 50O0 feet; free-stream Impact pressure, 10 to 22 inches water; cowl flaps, closed; thrust dish-loading coefficient, 0.09; propeller speed, 1200 rpti.

z > o >

z o

o

in

V»

Fig. 24 NACA TN No. I 109

CowlIng tealing ring

—Front row

—Rttr row

o Head preaaure, p A and C Hh2i + Hhat

B and D Hn4

E ph8

c Barral praaaura, p A and C Hb£|

B and 0 H|,4

E ph6

Tailad point» wara approx- imated


A BC DE Longitudinal pressure-tube location

Figur« 24. - Longitudinal pressure distribution tor closed cowl flaps. Density altitude, 6000 feet; free-atreai» Impact pressure, 23 Inches water; angle of attack of thrust axis, 2.7°; thrust disk-loading coefficient, 0.08; propeller speed, 1200 rpm.

N ACA TN No. I I 09 Fig. 25

Cowling sealing ring

Air flow

^L

Front row

—Rear row

o Head pressure, p A and C Hhgj» Hngt

B and D H„4

E ph8

D Barrel pressure, p A and C HD2i

B and 0 HB4

E ph8

Tailed points were approx- imated


A BC DE Longitudinal pressure-tube location

Figure 25. - Longitudinal pressure distribution for open cowl flaps. Densfty altitude, 5000 feet; free-strean I«pact pressure, 20 inches «ater; anale of stuck of thrust axis, 3.6°; thrust disk-loading coefficient, 0.1.0; propel ler speed, 1200 rpn.

Rennak, R. H. Uessing, li. E.

AuTKQ3(S) ktxa.

DIVISION, Power Plante, Reciprocating (6) SECTION: Cooling (1) CROSS REFERENCES. Enginee - Cooling teats (J2S62);

Thermal meaeuremeHt (93500)

iflfflO- 7*39 fixua. AGENCY Nu/aaa

REVISION

azn. TUUi Flight investigation of the cooling characteristice of a tao-roo radial engine installation - HI - Engine temperature distribution

^OSG-N-TITlEi

iTWG AGENCYi Hational Adviaory Committee for Aaronautics, Oashlngton, D. C. vnoN:

COUNWY U.S.

LANGUAGE Eng. r 'KC'NJTLASJj U.SJCLASS.

Onclass. DA« (PAGES IUÜST1

Oct'itol UP I 3g I ho tos, FEATUKS

tables, dlagrs,, graphs ADS7QAOT

Tro-rou radial engine was operated over aide range of conditions at density altitudee of 5000, ft and 20,000 ft. Quantitative results are presented ehouing effecte of flight and engine variablee upon average temperature and over-^11 temperature spread. Results indicate that temperature distribution patterns cere chiefly determined by fusl-air ratio and cooling-air distributions. Considerable change occurred in either spread or ehape of temperature patterns with variation of angla of attack of thru, t axis, everage fuel- air ratio, engine speed, poner, and blovar retio.

NOTE: Requests for copies of this report must be addressed tot H.A.C.A., Washington, D. C. .^

WHIGHT REU), OK». USAAF T-2, KS, AQ /AATEHia COMMAND An VECHXJCAI ÜNBZX

national advisory committee for aeronautics · 2011-05-15 · the cylinder numbering system used in...

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