wartime report - dtic · naca abb no. 5g20 national advisory committee fob aeronautics advance...

FILE COPY

-° !~W I ARRNo. 5G20

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WARTIME REPORT ORIGINALLY ISSUED

September 1945 as Advance Restricted Report 5G20

AN INVESTIGATION OF A THERMAL ICE-PREVENTION SYSTEM

FOR A C-46 CARGO AIRPLANE

VH - EFFECT OF THE THERMAL SYSTEM ON THE

WING-STRUCTURE STRESSES AS ESTABLISHED IN FLIGHT

By Alun R. Jones and Bernard A. Schlaff

Ames Aeronautical Laboratory Moffett Field, California

NACA WASHINGTON

NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not tech- nically edited. All have been reproduced without change in order to expedite general distribution.

A-48

J

NACA ABB No. 5G20

NATIONAL ADVISORY COMMITTEE FOB AERONAUTICS

ADVANCE RESTRICTED REPORT

AN INVESTIGATION 0? A THERMAL ICE-PREVENTION SYSTEM

FOR A C-46 CARGO AIRPLANE*

VII - EFFECT OF THE THERMAL SYSTEM ON THE

WING-STRUCTURE STRESSES AS ESTABLISHED IN FLIGHT

By Alun R. Jones and Bernard A. Schlaff

SUMMARY

As part of a comprehensive investigation of a thermal ice-prevention system for a Curtiss-Wright C-45 airplane, the change in stress at various locations in the wing outer panel caused by operation of the thermal system has "been determined in fli ght.

Wire resistance-type strain gages and thermocouples were installed at numerous locations in the wing. Recordings of change in stress and temperature resulting from operation of the wing-heating system were obtained for two speeds in level flight and one srjeed in a 2g bank.

Although the test results are directly applicable to the C-46 wing only, two general conclusions of interest in the design of wing heating systems are presetted. These are (l) the operation of a wing-leading-edge thermal ice-prevention system can result in stress changes which may (depending upon upon the unheated wing margins of safety) be negligible for regions aft of the double-skin region but will require investigation for the leading edge; and (2) local thermal stresses, possibly of critical magnitude, can be induced in sheet-stiffener combinations near the leading edge asxa result

NACA ARR No. 5G20

of temperature gradients in the order of 30° F between the stiffener and the adjacent skin.

INTRODUCTION

In the development of thermal ice-prevention systems for airplanes by the NACA, the principal objective has been to provide satisfactory and reliable systems for operation of the airplanes in icing conditions. Secondary problems, such as the effect of operation of the thermal systems on the airplane stresses, were afforded sufficient consideration to in- sure safe conduct of flight tests but were not examined in detail. The general acceptance of the practicability of thermal ice prevention for airplanes, however, has made necessary an investigation of these secondary problems in order to establish their magnitude and the degree of consideration to be assigned to them in future designs.

A preliminary examination of the effect of operation of a wing heating system upon the wing structure indicated that three factors were of importance: namely, (l) a reduction in strength and elasticity of the wing material caused by elevated temperatures, (2) increases in stress in the heated leading edge because of restrained thermal expansion, and (3) increases in stresses in the remainder of the wing, which is relatively unheated, caused by expansion of the leading-edge region. Although the first two items listed were the most ob- vious on first consideration of the problem, the incipient failure of a Lockheed 12A wing outer panel at the rear spar due to excessive expansion of the leading edge, as described in reference 1, was evidence that the last item might prove to be the moBt critical of the three.

The purpose of the investigation reported herein was to measure the changes in stress in a typical wing which result from the operation of a thermal ice-prevention system in order to establish the magnitude and seriousness of these changes in stress upon the structural integrity of the wing. The testB were conducted as the seventh part of a comprehensive investigation of a thermal ice-prevention system for a 0-46 airplane. The first six parts of the investigation are presented as references 2 to 7.

The flight tests were conducted at the Ames Aeronautical Laboratory of the National Advisory Committee for Aeronautics, Moffett Field, Calif., at the request of, and in cooperation with, the Air Technical Service Command of the U. S. Army Air

NACA ARR No. 5Ö20

forces. The authors wish to acknowledge with appreciation the valuable assistance, during the. conduct—o-f- the flight tests, of Mr. Robert Deland of the Curtiss-Wright Corporation

Description of Equipment

The C-46 airplane'(fig. l) is a twin-engine, low-wing, transport-cargo monoplane powered by Pratt <5-, Whitney model R-2800-51 engines having a sea-level rating of 200C horse- power each.' The wiflg' span is 108 feet, the wing area 1360 square feet, and the normal gross weight 45,000 pounds.

The details of construction of the left-wing outer panel in which the stress measurements were obtained are shown in figure 2. The wing is of all-metal stressed-skin construction with spars -at 30 and 70 percent of the chord. The skin is reinforced with spanwise hat sections and extruded stringers. At the wing root (station 0) the outer panel is attached to the inboard section through splice angles along the upper and lower surfaces of the wing. The wing profile varies from an NACA 2301? section at station 0 (chord = 198 in.) to an NACA 4410,5 section at statiöri 412 (chord = 66 in.).

A complete description of the revisions to the C-4 6 airplane for thermal ice prevention is given in reference 4; however, a brief outline of the alterations as- they affect the wing structure will also be given herein. The heated air for ice prevention is obtained from exhaust-gas-to-air heat exchangers installed on each side of the nacelles as may be seen in figure 1. The air from the outboard exchangers is directed to the outer-panel leading-edge heating system shewn in figures 2 and 3. The details shown in figure 3 comprise the major revisions to the Wing, the only other alterations being small reinforced holes iff the spar webs (fig. 2) for circula- tion of the heated air. '

The changes in stress were measured With standard Baldwin- Southwark wire resistance-type strain gages. In determining the location of the gages, consideration was given to the stress reports for the unheated wing and the measured temperature distribution in the wing as presented in references 5 and 7. The stress reports showed that the margins of safety were appreciably lower at the wing root than near the tip; while the temperature' data indicated that, in general, the spanwise distribution of temperature rise for a given chord location was approximately constant'. The assumption was made that, for a given chord location, the stress change along the span would

NACA ARR No. 5G20

be substantially constant with the application of heat; and, therefore, the gages were located at two stations relatively- near the root (47 and 137 in1, from station 0).

The individual locations of the strain gages and thermocouples are shown in figures 4 and 5. Two general types of gages were employed: namely, single-element gages on structural members such as hat sections and spar caps where the direction of stress was known, and triple-element gages on the wing skin where the direction, as well as the magnitude, of the maximum stress had to be established. All the single- element gages (hereinafter referred to as plain gages) are designated in figures 4 and 5 by the,letter G before the gage number and were Baldwin-Southwark type A-7 gages. The triple-element gages (hereinafter referred to as rosette gages) are'designated in figures 4 and 5 by the letters RG, and were Baldwin-Southwark type.ABR-4 gages forward of 10- percent chord and type AR-1 gages aft of that point. The ABR- 4 gage is a high-temperature type required on the heated leading edge. The discontinuities in the strain-gage numbering system are the result' of omitting from the list of gages those which failed.

The plain gages were- installed with their strain-sensitive axes parallel to the longitudinal axes of the structural members. The rosette gages were oriented as shown in figure 6, which also gives the designation for the three strain elements. The fact that the high-temperature rosettes were supplied in a delta, arrangement and the low-t emperature rosettes in a 45° pattern has no particular significance. Aft of the double- skin region the rosette gages, were mounted in pairs, back to back, and connected in series in order to eliminate strain indications caused by possible local skin buckling. This procedure was not feasible in the double-?skin region but was considered unnecessary because of the stabilizing effect of the inner corrugations. Two typical plain-gage and one rosette- gage installations are shown in figure 7.

The strain-gage recording equipment is shown installed in the airplane in figure 8 and consisted basically of a 13-chan-

t] b< gages at one time to the recording, oscillograph... A schematic diagram showing the complete- circuit for one chj sented in figure 9. The active and dummy gages wOJ.o ^uunovooi in two arms of ä Wheatstone bridge circuit, the remaining two arms consisting of variable resistors in the four-channel

! #

NAOA ARE No. 5G20 5

bridge boxes. Any gage could be connected into any channel of the oscillograph by plugging it into the proper outlets of the strain-gage switch box. The change in resistance of the copper strain-gage lead wire caused by elevated temperatures in the wing was considered to be of sufficient magnitude to require the use of a compensation lead wire installed adjacent to the active lead wire of each gage. Means for calibration of all channels of the recording oscillograph at any time during the test program was provided by the installation of a variable-resistor switch and. a channel-selector switch in parallel with the dummy gages.

Surface-type i r on-constantan thermocouples., rolled to a thickness of 0.002 inch, were cemented to the aluminum surface with* the junctions within one-quarter of an inch of the strain-gage elements. The thermocouples are designated by the same number as the corresponding strain gage and are pre- fixed with the letter T. (See figs. 4 and 5.) A thermocouple designation shown in figures 4 and 5 which is not ac- companied by a corresponding gage number indicates the gage failed. In the case of the rosette gages aft of the double- skin region, where two gages were installed at each location, a thermocouple was placed on each side of the skin in the event that an appreciable temperature gradient might occur. The temperature readings were recorded with a Brown recording self-balancing potentiometer shown in figure 8. Additional thermocouples which had previously been used in the performance-test installation (reference 5) were also connected to the Brown recorder. These additional thermocouples permitted the measurement of the ambient-air temperature, the air temperature at the heat-exchanger outlet, and the air and skin temperatures in the double-skin region at stations 159 and 380.

Standard NACA instruments were installed to record airspeed and normal acceleration. An NACA timer, shown installed at the left of the airspeed recorder in figure.8, was used to synchronize the airspeed, accelerometer, and oscillograph records. The accelerometer was installed on the floor of the airplane at the center of gravity and oriented to record accel- erations normal to the wing chord at station 0. The airspeed recorder was connected to the service Kollsman airspeed-head installation and the error in static-pressure reading was determined in flight for all the test conditions with a trailing static pressure head.

ZTJLCA ARB No. 5&20

TEST PROCEDURE

Instrument Technique

When employing strain gages to measure stress changes in a structural mem"ber which is subjected to a temperature, as well as a stress variation, a question arises concerning the differentiation between movement due to stress and that due to..thermal expansion. Superimposed on thi s complicat i on is the unknown effect of elevated temperatures on the resistance of the strain-gage material. One installation method common- ly used to overcome these problems consists of locating the dummy strain gage beside the active gage but cemented' to a small piece of the structure metal which is free to expand. The compensation if based on the assumption that the piece of metal to which the. dummy gage is attached will assume the same temperature as the surface upon which the active gsge is installed. This method was considered unreliable in the case of the C-46 wing.

Laboratory calibrations were made to investigate the stress-strain curves which would result if the dummy gage ware maintained at a constant temperature while the aluminum calibration specimen and the active gage were subjected to various stresses at different temperatures. The results of this calibration for a typical strain gage used in the investigation are- presented in figure 10. The important fact to note in figure 10 is that the calibration curves form a series of parallel straight lines. This means that the sensi- tivity factor of the gage (ratio of unit change in resistance to unit strain causing this change) is independent of the initial stress and temperature throughout the calibration range.

The method of applying the calibration curves of figure 10 to the determination of stress changes in the heated wing was as follows: Assume that point a of figure 10 represents the stress and temperature conditions at a gage in the wing before the wing heat was applied. The temperature of point a was known but the stress -was not; however, since only stress changes were to be evaluated, the strain-recording equipment was adjusted to zero reading (or balanced). The heat was then directed to the wing and the changes in temperature and strain-gage reading were recorded. If the temperature change had been 30° F and the strain-gage reading agreed with point b, pure expansion and no change in stress would be indicated. If the strain-gage reading corresponded to point c, however, a change in stress equal to the distance between points c

«

NACA ARE No. 5G20

and d '(or an increase in tension) was indicated. It is evident that the same result would he ohtained for any other initial point of e instead of a since the curve slopes were all equal. Hence the absolute value of the stress at the time of "balancing the strain equipment had no bearing, on the result .

Plight Procedure

All the flight tests were made at a pressure altitude i

of 10,000 feet and a take-off gross weight of 45,700 pounds. Data were obtained at two airspeeds (110 and 135 mph., indicated) in level flight and at 155 to 165 miles per hour in a 2g bank at constant altitude. The airspeed in the 3g bank was calculated to give the same lift coefficient as that ex-

D

isting in the level-flight condition at 110 miles per hour in order that the two wing-loading conditions would be directly comparable with respect to chordwise pressure distribution.

In the case of the level-flight tests, the airplane was stabilized in flight at the desired speed and with the wing- heating system in the off condition. Twelve gages were plugged into the'strain-gage switch box, the equipment was balanced, and a calibration record was obtained. A 2-minute record was then taken with all the recording instruments. The length of the record was established by the time required for the thermocouple recorder to complete one cycle. At the completion of the heat-off recording, the heated air was directed to the left-wing outer panel. Preliminary tests indicated that 4 minutes was an adequate period of time for sta- bilization of the wing temperatures. At the end of 4 minutes, therefore, a second 2-minute record was taken with all recording instruments. This procedure, the heat-off recording and heat-on recording, was then repeated with 12 new gages until the entire gage installation had been included. The stress changes due to wing heating were taken as the difference between the stress records for the two flight conditions just outlined with corrections being made for temperature effect.

In the case of the 2g banks, a slightly different procedure was followed because of the difficulties involved in maintaining a reasonably constant acceleration over the entire 2-minute period required by the thermal recorder. The airplane was flown at an indicated airspeed of 110 miles per hour in a level attitude with the heat off, and the strain-gage equipment was balanced. Records were obtained for this condition with all recording equipment. The airplane was then banked to produce a 2g normal acceleration at an indicated

NACA ARR No. 5020

airspeed of 155 to 165 miles per hour and constant altitude. At a signal from the pilot indicating stähle flight condi-' tiohs, a 1-mlnute record was obtained. The airplane was re- turned to the level attitude (thermal system still heat off), the airspeed was set at 110 miles per hour, the strain-gage equipment was rebalanced, and complete recordings were taken. The engine uower was then increased to that required in the banked condition (appro*. 32 in. Hg manifold pressure and 2200 rpm, engine speed), the thermal system switched to heat on, and the wings allowed to heat for 4 minutes in level flight. A complete temperature record was taken at the end of this period, and the airplane was then banked to the 2g attitude. Again at a signal from the pilot, a 1-minute record of all recording instruments was obtained. Although this 1- minute period was adequate for recording the stress records, only half the temperature data could be recorded in this in- terval. The temperature readings in banked flight, however, agreed within the limits of experimental error with the temperatures recorded immediately: before the bank. Eence the level-flight-temperature data were considered satisfactory for the banked condition.

The differences between the stress records for level flight at 110 miles per hour, heat off, and banked flight, heat off, supplied the streBS changes due to increased wing loading with no Increase in temperature. The differences between the stresB records for level flight at 110 miles per hour, heat off, and banked flight, heat on, supplied the stress changes due to increased wing loading plus those due to wing heating. The stresses due to wing heating alone were determined by subtracting the stresses due to increased wing loading from those due to increased wing loading plus heating.

PRECISION OF DATA

An exact determination of the accuracy of measurement of the stress changes was not possible because the evaluation of some of the factors involved was not practicable. A laboratory check of the precision of the entire strain-gage circuit and the flight-test procedure, however, presented some indication of the degree of accuracy of the final results. This laboratory check was undertaken with the equipment shown in figure 11. A strain gage and a thermocuple, each identi- cal to those installed in the C-46 wing, were installed on a tension specimen of aluminum alloy 24S-T alclad. The specimen was subjected to various known stresses at different

.'

NACA ARS No. 5020

temperatures (heating being obtained with the lamps shown in fig. 11) and the stress and temperature changes recorded with the flight-test equipment.

These laboratory records were evaluated using the same calibration curves employed in establishing the flight-test results and were compared to the known stress changes. A precision of measurement of ±200 psi was established. Adding to this the unknown errors in constancy of flight conditions, effects of vibration on instruments, and so forth, an accuracy of ±400 psi for the stress-rchange data presented herein is believed to be.reasonable.

Based on laboratory calibrations of the iron-constantan thermocouple wire and the Brown recording potentiometer, plus an unknown installation error, the temperatures presented are considered accurate to ±3° J for the strain-gage locations and ±5° P for the remaining installations. The 25° P correction for surface thermocouples forward of the baffle (5 percent chord) mentioned in reference 5 has been applied to the data presented for stations 159 and 380.

The airspeed data are considered accurate to ±1.5 miles per hour for the level-flight tests and to ±3 miles per hour for the banked condition. The error in recorded acceleration may be taken as ±0.02g.

RESULTS

The test conditions and the resulting total heat flow to the left wing are given in table I. The stress changes measured by each gage element are presented in table II, part 1, for the plain gages and in table II, part 2, for.the rosette gages. The data presented in table II have been corrected for temperature effect by the method previously outlined in the discussion of flight procedure and figure 10, and represent actual changes in stress for the plain gages. For the rosette-gage data (pt. 2 of table II) a further correction based on the Poisson ratio effect on the gage reading's must be applied.

The maximum and minimum changes in normal stress, the maximum change in shear, and the direction of action of these stress changes were computed from the rosette-gage .data (table II, pt. 2) and are presented in table III. The terms "maximum normal" and "minimum normal" are used here in their

10 NAOA ARB No. 5&20

algebraic sense. The PoisBon ratio correction has been applied, and the data presented in table III represent actual; stress changes. The angle Q in the table gives the direction of the line of action of the maximum change in normal stress and 1B measured from the wing chord with the positive direction as shown in figure 6. The line of action of the mininmm change in normal stress is at right angles to that of the maximum normal stress, and the maximum change in shear is at an angle of 45° to either of these axes. The rosette-gage stresses were also resolved in the spanwise and chordwise di- rections and these data are presented in table IV.

DISCUSSION

The anticipated effect of heat-in* the wing leading edgß on the chordwise wing «tress distribution was (l) an increase in compression for the double-skin region as a result of resistance to thermal expansion, (2) an increase in tension in the region between the double skin and the 30-percent spar caused by the expanded leading edge pulling on the relatively cool afterbody, and (3) an increase in compression at the 70- percent spar caused by the two spars and the wing skin acting as a box beam to resist the moment imposed, by the expanding leading edge; This general trend is evident in all the curves showing the chordwise distribution of stress change. (See fig. 12.) These curves are based on the values of stresB change normal to the.chord presented in tables II and IV. The chordwise temperature distribution has been added to the stress-distribution curves in order to facilitate the inter- pretation and explanation of the test data.

A comparison of the stress changes for the three flight conditions is presented in figures 12(g) and 12(h). Although the increased heat supplied to the wing in the 2g bank is evidenced by increased,compression at the leading edge and some deviation of the 2g curve from the level-flight curves at other chord positions; the over-all agreement between the stresses for the three conditions is considered sufficient to allow them to be discussed as one general trend.

NACA AEE No. 5&20 11

The expected compression at the Reading edge and' 70-percent spar is evident in figure .12. Ir'om ,10-to 30-percent chord, however, considerable variation in the data is noted. This apparent discrepancy at, first appears to refute the anticipated general trend, "but on further examination is seen to be the result of local conditions superimposed upon the over-all pattern. An example of this effect is- shown by the data presented for gages 41, 51, and 52 in figure 12(d). The expansion of the leading edge would be expected-to pull on the hat section containing gage 51 and exert tension similar to gage 36 on the lower surface. Apparently, however, the heated air discharging from'the double skin heated the hat section considerably (note temperature distribution) and the restrained expansion induced couipressive stresses which were • larger than the induced tension. The expansion forces of the hat section, in turn, placed the colder skin in tension as signified by the indication of gage 41. At the location of gage 52' the ' stress has again reversed, the actual, value at gage 52 being the result of the combined effects of several factors of unknown magnitude.

Another interesting example of large stress changes caused by local- conditions (temperature gradient between stiffener and skin) is presented by the indications of gages 50 and 60 in figure 12(d). Although the two gages are located within 1 inch of each other, the stress in the angle was about 3000-psi compression; whereas the skin was practically un- stressed. An examination of the t emperatur e.-di stribut i on curve shows that the - angle temperature.was approximately 10 F greater than the skin temperature. For a•completely restricted aluminum member the increase in compressive stress for a 10° F temperature ' rise is about 1300 psi . It appears reasonable, therefore, to picture the stress changes occurring at gages 50 and 60 as a uniform increase in stress in both the angle and skin until a value of about 1500-psi compression has been achieved. This is followed .by an increase in compressive stress in the angle (caused by an increase in temperature of the angle only) and a decrease in stress in the skin because of the stretching action of the angle.

Because local conditions in some cases caused large stress' differences at a given chord location (the actual deviation depending upon whether the strain gage was mounted on the, skin or on a longitudinal stiffener) the stress curves presented for station 137 represent some mean values of stress for the region from 10-* to 30-percent chord. Although the same local heating existed at station 47 (note temperature distribution for upper surface, figs. 12(a), 12(b), and 12(c)), the scatter of data was not obtained because all the gages in the

12 NACA ARE No. 5&20

local heating region were mounted on the stiffeners. The stress curve for the upper surface is, therefore, more repre- sentative of. .str.esB in the stiffeners than in the skin.

The belief that the scatter of data between 10-and 30- percent chord ie-largely attributable to localised heating , rather than inaccurate measurements is verified by an inspection of figures 12(g) and 12(h). These curves show consist- ency of the data between the three flight conditions for the regions: where, temperature gradients between the internal structure and skin were negligible. Further examples of this local heating effect between 10- and 30-percent chord could be cited; .however j the purpose of this report is to determine, in generali the magnitude and importance of the stress changes rather than to present a detailed investigation of the 0-46 - wing..

In order-to obtain some indication of the seriousness of the stress changes, the test data were compared with the critical values as specified in the wing outer~panel stress analysis as prepared by the airplane manufacturer. The flight conditions considered critical for the wing outer panel in the stress analysis were conditions I, II, Uli, and VII as presented in reference 6. The assumption is made that the thermal system might be in operation during any of these flight conditions; and, therefore, the measured stress changes are compared with (l) the allowable stress and (2) the critical margins of safety. A complete and precise comparison was not practicable because (l) some of the strain-gage locations were not considered critical in the stress analysis and, therefore, no margin of safety was presented, (2) the nearest station to 137 covered by the stress analysis was station 112 (or 25 in. inboard of 137), and (3) in the leading-edge region, the type of wing construction was considerably different from that uBed in the stress analysis.

The stress changes in the longitudinal stiffeners and spar caps ranged from 1 to 5 percent of the allowable stress with the exception of a few stiffeners near the leading edge. Whether or not such changes are critical depends, of course, upon the particular margin of safety for each stiffener. For the specific case of the C-46 wing the margins were sufficient- ly high in most cases to absorb an increase of 5 percent of the allowable. It would be inadvisable, however, to make the general statement that increases in stress of this magnitude are not critical because of the specific nature of the problem. For example, the increase in stress of the T-section containing gage 52 at station 1.37 was only 8.00-psi compression, or

NACA ARE No. 5G20 13

about 2 percent of the allowable stress for that member. This increase, however, was sufficient to reduce the already critical margin of safety from 0.06 to 0.03. The spar caps, on the other hand, because of their large margins of safety may be considered as unaffected by the thermal stress changes.

In the case of longitudinal stiffeners located near the point of discharge of the heated air from the double-skin region (gages 50 and 51),the changes in stress ranged from. 10 to 16 percent of the allowable stress. The- value of IS percent was obtained with gage 50 and the increased stress (2800-psi compression) would be sufficient to change the margin of safety for the angle in compression from 0.08 to -0.08. Although this negative margin is small the important fact to note is that impingement of the heated air directly upon a structural member may induce stresses which are large enough, to change the margin of safety from a satisfactory to an un- acceptable value. In this connection it is of interest to note the effect of stringer configuration on the temperature distribution. An inspection of the upper-surface temperatures for either station presented in figure 12 shows that a larger temperature gradient between the skin and the adjacent stiffeners, and hence larger local streBB changes, was measured for the hat sections than for the T-sections. In regions where heated-air discharge is apt. to cause local overheating of the wing structure, therefore, the air should be directed away from the structure by nonstructural vanes and, if practicable, the.structural member should have a maximum of area in contact with the skin.

The maximum shear-stress changes in the skin aft of the double-skin region, as measured by rosette gages 4, 32, 41, and 60 and presented in table III, were all of low magnitude (about 2 percent of the allowable shear stress). The margins of safety at these locations were all large and quite capable of absorbing the 2-percent increase in stress. As mentioned previously, however, this general conclusion should not be applied indiscriminately to other wing structures.

The largest changes in stress were measured at the wing leading edge,but are particularly, difficult to interpret because of the lack of data on the allowable leading-edge stresses. Some test data are available which can be reasonably applied to the C-46 leading-edge construction, prior to revisions to incorporate the thermal system, but information on the double-skin type of construction did not appear to be available. In order to obtain some indication of the seriousness of the measured leading-edge stress changes, the revised

14 FACA ARR No. 5G20

wing stresses will Toe compared to those allowable for the. un- altered leading edge. The;structural effect of the inner corrugations will of necessity remain an unknown factor in this report.

In the wing structural analysis for the unrevised leading edge the skin was considered to carry only shear.due to torsion and beam bending. A stress element on the leading edge at station 47 (location of gage 7) would therefore be repre- sented as shown in figure 13(a).. The chordwiee shear stress of 7300 psi represents the maximum value investigated in the stress analysis, and compr©S6ion.in the leading edge due to chord bending was considered negligible; Operation of the thermal system produced additional normal stresses of approximately 4000-psi spanwise compression, and 5000~psi chordwise tension (data for gage 7i table IV), as shown in figure 13(b). Since these normal stresses can be considered the principal stresses (table III, angle 6, approx. 0°), no additional shear is added to the stress element in a chordwise direction. The actual stresses existing at gage 7 during operation of the thermal system, therefore, may be taken as shown in figure 13(b).

The allowable stress for the unheated leading edge was presented in the stress analysis as the stress at failure in an unstiffened thin-walled .cylinder in pure torsion. (See reference 9.) .From reference 10, the-heated leading edge can be approximated by a cylinder in combined loading for which the three allowable stresses are related by the equation as follows:

££)' + '-F

•p ^ b ,* . c

8

where

fc allowable compressive stress in co'mbined loading

Pc allowable compressive stress for pure compression

f8 allowable shear stress in combined loading

F8 allowable shear stress for pure shear

ft allowable tension stress in combined loading

F^. allowable tension stress for pure tension

a,b,c empirical exponents

NTACA ARE No. 5G20 15

No data were available on tests of cylinders in combined compression, tension, and shear from which exponents a, b, and c could be evaluated. If the skin is assumed capable of de- veloping the ultimate tension stress in pure tension (i.e., if Ft = 56,000 p8i), the factor ft/^t may te neglected. A consideration of references 10 and 11 relative to the aerodynamic shear stresses and the thermal compressive stresses indicated that equation (l) could be expressed as

k • (k) -i <2)

The allowable shear stress due to pure torsion Fe may be taken as the same value presented for the unheated leading edge or 9700 psi. This value was obtained from reference 9 for values of l/r = 2.3? and r/t = 158

where

I length of cylinder

r radius of cylinder

t skin thickness

In a like manner the value of $c is taken from reference 12 (curve c of fig. 7) to be 18,000 psi. In a rigorous analysis these allowable stresses would be reduced by about 10 percent because of the effect of elevated temperatures on the tangent modulus of elasticity. (See reference 13.) This factor has been neglected, however, because of the general nature of the discussion. If the torsional and bending shear due to aerodynamic loading is assumed to remain constant, then fs = 7300 psi. (See fig. 13.) Substitution of the foregoing values in equation (2) provides an allowable compression stress in combined loading of fc = 7800 psi.

A comparison of the actual compression stress (4000 psi) with the allowable stress (7800 pei) indicates that the thermal system has not resulted in a critical stress condition. Such a statement, however, should be qualified with a few re- marks which are of interest. The compression load in the leading edge due to chord bending, although negligible for the unrevised wing, cannot be neglected for the heated wing. The wing structural analysis indicates that the compression

16 HACA ARE No. 5G20

due to chord "bending is.at»out 5000 pel,which, when added to the thermal compression, gives a total value of 9000 psi. This, value is 1200 psi in excess of the allowable. It is not intended that these figures should be taken as a precise In- dication that the allowable has been exceeded, because of the uncertanties Involved in equation (2) and the unknown increase in allowable stress afforded by the inner corrugated skin. It is intended, however, to point out that the thermal stresses superimposed on the aerodynamic stresses might produce a critical condition in combined loading, particularly in cases where the inner skin cannot appreciably increase the allowable stress of the outer skin.

The effect of elevated temperatures on the strength and elasticity of the wing material is a subject of considerable interest in the design of thermal ice-prevention equipment, but a detailed discussion is regarded as beyond the scope of this report. The generally accepted temperature limit for aluminum-alloy structural members is 200° F. Above this temperature, reductions in the allowable stresses are required. An examination of table V indicates that certain leading-edge components such as nose ribs may exceed this critical temperature, resulting in a lowering of the allowable stress, but structure aft of the leading edge can be maintained at sub- critical temperatures. Although researches sUch as that presented in reference 13 are providing valuable information on the strength of aircraft structural materials at elevated temperatures, further investigations appear desirable to establish the effect of cyclic heating over a long period of time on the strength of aircraft materials and to examine the phenomena associated with metal "creep" at increased temperatures. The extension of airplane speeds into the sonic range Will result in aerodynamic heating of the airplane surfaces of appreciable magnitude, and thermal stress problems will be

NACA ARR No. 5G20 17

unavoidable. Thus additional research on wing thermal stresses, particularly in the leading-edge region, will be of value in the development of wings for high-speed airplanes as well as in the development of thermal ice-prevention systems.

CONCLUSIONS

From flight tests of the change in stress in Wing structure resulting from the operation of the thermal iCe-prevention system, the following conclusions are made:

1. The operation of a wing-leading-edge thermal ice- prevention system can result in stress changes which may (depending upon the unheated wing margins of safety) be negligible for regions aft of the double-skin region but will require investigation for the leading-edge region.

2. Local thermal stresses, possibly of critical magnitude, can be induced inn sheet-stiffener combinations near the leading edge as a result of temperature gradients in the order of 30 F between the stiffener and the adjacent skin.

Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Moffett Field, Calif.

REFERENCES

1. Rodert, Lewis A,, and Clousing, Lawrence A.: A Flight Investigation of the Thermal Properties of an Exhaust- Heated-Wing De-Icing System on a Lockheed 12-A Airplane NACA ARR, Suppl. no. 2, Sept. 1941.

2. Neel, Carr B., Jr.: An Investigation of a Thermal Ice- Prevention System for a C-46 Cargo Airplane. I - Analysis of the Thermal Design for Wings, Empennage, and Windshield. NACA ARR No. 5A03, 1945.

3. Jackson, Richard: An Investigation of a Thermal Ice- Prevention System for a C-46 Cargo Airplane. II - The Design, Construction, and Preliminary Tests of the Exhaust-Air Heat Exchanger. NACA ARR No. 5A03a, 1945.

18 NACA AHB No. 5(JC0

4. Jones, Aluo, and Spies, Ray J., Jr.:,. An Investigation of a Thermal Ice-Prevention System for a C-46 Cargo Airplane. , III - Description of' Thermal Ice-Prevention. Equipment for Wings, Empennage, and Windshield. NACA ARR No. 5A03b, 1945.

5. Selna, James, Neel, Oarr B., Jr., and .Zeiller, E. Lewis: An Investigation of a Thermal Ice-Prevention System for a 0-46 Cargo Airplane. IV - Results of flight Tests in Dry-Air and Natural-Icing Conditions. NACA ARE No.. 5A03c, 1945.

6. Selna, James: An Investigation of a Thermal Ice-Prevention System for a C-46 Cargo. Airplane. V _ Effect of Thermal System on Airplane Cruise Performance. NACA APR No. 5D06., 1945. .- .

7. Selna, James, and Kees, Harold L.: An Investigation of a Thermal Ice-Prevention System for a C-46 Cargo Airplane. VI - Dry Air. Performance .of Thermal System at Several Twin- and Single-Engine Operating Conditions at Various. Altitudes. NACA ARB. No., 5C20, 1945.

8. Anon: Airplane Airworthiness. Civil Aeronautics Manual 04, Civil Aeronautics Administration, Feb. 1941.

9. Lundquist, Eugene E.: . Strength Tests, ,on Thin-Walled Duralumin Cylinders in Torsion. NACA TN No, 427, 1.932.

10. Anon: Strength of Aircraft Elements. ANC-5, Army-Navy- Civil Committee on Aircraft Design Criteria, Dec. 1942.

11. Guggenheim Aeronautical Laboratory, California Institute of Technology:, Some Investigations of the General Instability of Stiffened Metal Cylinders. VII - . Stiffened Metal Cylinders Subjected to Combined Bending and Torsion. NACA TN No. 911, 1943.

12. Lundquist,. Eugene E . : Strength Tests of Thin-Walled Duralumin Cylinders in Compression. NACA Rep.. No. 473, .. 1933.

13. Flanigan, A. E., Tedsen, L. F., and Dorn', J. E.: Final Report on .Study .of the Forming Properties of Aluminum Alloy Sheet at Elevated Temperatures. Part X - Tensile Properties after Prolonged Times at Temperature. Serial No. W-146, NDRC, Oct. 20, 1944.

NACA.ARR No. 5G20 19

g

jnoq aed n^g 'Sup» o* moxj ^BBH

anoq J9d BDtmod *3up» o* motj ajy

WM

&0 'aöSuBqox© ^»oq jo ^9i^no 4« 9.xri4Bjedui93 apf

Xjnoaew jo 89U.OUT '9ans89Jd PIOJIUVH

tad a *p99de BUfSug

o c o

CM

UN CO

ITS UN HN

tf

O o ON

3o~ *8an^»Jt9dm95 <xp»-4u«TqnY

*99j '•pn^j^x« •ansB9J<i

epunod *jjo-93f»$ S» ^qSy9Ä BSOJtfl

c o

T3 C O O

P

•a

o o o o H

O O

UN

00

p o • p

fe-ä O «H OH

O-H

O «

o o o CM CO rvj

o CM o

UN

CM

O o ON

o o o » o

O O

UN

to < M

P O 0) (H

O O

•a

-EH

UN • KNH

O o o « o o-

O

ON

UN

CM

O O CM CM

O O O

O O

UN

•E

p o o (H U O o

. . m

UN 60 VOCM H

• O 03

*N UN • UNH H

30 NACA ARR No. 5G30

38

H^

o O O o O O r-l cr UN o CO ON 3 ON KN US t- ON UN OJ -d- CiJ 1 1 i o H

1 H

1 H

1

o o o o O O o H OO c- c- -=*• CO vo *o c- c— c- UN H CM UN Ü 1 1 i O H H H

1 1 1

r- o o o ON O O o OJ CO UN c— £ H KN -=* tt ON ON ON KN KN UN + + + + + + vo o o o UN O O o Si KN

UN CO UN

UN KN

UN O 3" -3- + + + + + +

UN o o O -Cj- o o o 3 VO o c- UN o vo c~ + H -=*• tt H H CM

1 1 + 1 1

o o o o O o

1 UN UN UN C\J ON CO H -3-

1 vo

1 vO H

1

UN i

+ co 1

5 o o o o o CM o e\j H c- KN UN

ry co co i UN o CM vo e> + • CJ <\J CM CM

-3 KN H

1 1 1

o o o o O o ON CO CO CO o o O -3

fi d -3- UN UN a £ UN H KN o O H H • o CM CM CM

P od

•f + «H P cd

1 1 1

o o o O o P GO -st- KN o p co H ON ON 00 H o r-l co 00

tf vo KN UN

o r-l H UN H H + + • • + + o Ö " 6 o O o

c- o H H r- CO t- KN

Ö H

f CO + H + *

CVJ H

o H +

VO +

o o O o O O UN t- 5 UN UN CM UN ON c« c— t*- -d- UN t- ON o + +• + . O 1 1 H

1

• • P • • P oo 00 O w 00 o

• • © • • « •< < fc << << u

• / • • u • / • • u to / M H o to / M M o d / o « / o to P P «•a to / P P

** O O o O O P © P a © «DP ©P 0. 03

6ft 6ft S.Q 6ft 6ft BP O-H O «H UN t0 0«H 0«H UN t0

/ ö OH OH ^OCM / ö OH OH VOCM / o «* «H H / o *-i <H H / *"-t A £ • / *H .G A • /P P P.H &H o oo /P P an O.H O 00

/«•O B£ B£ P •

<< /ftJo BS B2 p •

< / -H C O 0 UN 4) UN • -H C o © UN© UN •

1 Eo i-IH H

MNH H

UNM H /Co H

KNH H

UNM H

A o a • •H A o © C fc

*H © P

© O*

ä a

2 u o* © M a

(H « © *d a § "0 a § «* o © a ©

© * u

© P a © © fc a P O a •H

m fi © o ©

*H U a a a 6 © o p Ü

8 s p P o o c fi © c «o «a

> I

Oft

+ I

P

HACA ARR No. 5G20 31

£! O-WH

CM

.»•""•t

o O

O o O O t- o O o o CM

UN %? i H H

ON H

KN r* *r NO

NO s CM KN CM K\ CM + 4

1 1 1 • 1 1 a o O O ON o O O

O H o CD

s M CM ON CM •£> •• p H c- P <t ON -* «•« H CO CO o V. -* f«N. Si (d •H -* 1

as CM H CM « CM K\ p 1 1 1 1 1 1 1 M

o o O o O o o o •4 t>- 00 IT» H 00 ITN o O w c- o

«-I co NO ON u 00 CO CO NO *H ^

CO UN + UN

4 "? -f + •H 4

m o o s O

UN m o CM

O O

o UN s

r-l . *

o CO

o o

u H CM

CM CM

vo CM

vT 3 UN -4" r-

«H UN 4

KN 4

1 1 i 1 1 1 KN c t>- H O

l*- o o o KN CO o O o v< o O ^t c^ * in 00 o H !£> N CM o -=f •K) P .M UN r-

o *7 -3- UN CM O «*T H «c C7\ s « «>r o ON a PS KN K\ UN a 03 IT> -d- UN P <Ü H o I 1 i o 1 1 1 10 4 4

*r4 •P o o o p o o o

UN 1

O

-p a 8 o

NO 00

• p *r KN

IT\ U>

CM O UN

UN

UN • q 3 «r 3

• + • 4 + 4 4 o PS P « o o o

P O O O P <^** o o

N 00 H N> M IfN O 00 5 •?"•

» .w UN ON

« »7 t— UN -=f a vT IT» IT» H *r i-H 1 0 CM CM KN 4> CM H H a 1 P P

1 • | P P

1 1 1 P O

• o o o e t- O O O a «-I Q o «0 O tf

IT*

8 H CO

UN NO o & •? CO

H £ CO ON «

•p ctf *? S UN

UN t, 03 H H »4 as CM H 1 • P 1 1

m 1 1 1

« 1 t P

P a

p o o O P o O o • O o H M 00 SO C- H «4 CO K\

3 « CM <« o KN

9 *7 5 % 3 •a «M IfN

NO <3 O © *i UN 4

UN 4 + 4 4 + + • OS

•

P i? o o -3-

o 8 CO

•

*r o

ON

o UN CM

£ "p *? o 1

m UN UN -=f a H 1 + • -=t 4

« + 4 4 Ss 1 t c o, o o o a. NO o

& O O O

a £ • t- CM KN S K> «i KN 4 •rt « UN H 5 %7 CM K\ CO S O *7 CM CM h P *7 KN O •P OS KN S CM P 03 KN KN • « H H

1

4 • 4 i

•ft + •f

• P • 4 4

O

o O o O O O O

«rl w IM 8

ON

8 •H n

•4

3 NO

CM

P 1 » tf

«? ON CO KN 4 + + • • + 3 os

1 • • P • * P » « 09 01 o 0) 60 o 01 01 • • • • • • j • •

< ^ u •< «« b / ^ < V / * • u • / » • fc. • / • • to / M H o 60 / M H o 60 / H H « / o a) / o * / to / P

O • P

P P • P as

60 / P O

p

gp *1 60 / P O • P

P O • P

64 6« B» 64 64 B^J 1 64 e-fl OvH 0<H UN 61 O «rt o*-i UN 60 \ 0«H o«-t

/ ö OH OH NO CM / o OH OH ^OCM l ö OH °d / o «H «4 H 1 © «H «H »-« / o _*« ÄH *H

• O 0} *P P fn S^ •

OOl 1*3 4» fl.H t« Äf B? B£ p •

•4 AftS ss s5 p • /•äS fl* *%

O • ir\« UN • /H Ö O 9 UN« UN • 1 * a O • UN«

f El» H KNH H

tfNH H / E o H H KNH

H UNM H

/ C* O HH H

KNH H

• c

S.1

« «

• • •o-o

4 I

P

p

p o

P 0

a o *

£ u

• p o • 6 o

i « p

p


I EH EH

• • H 73

V f l-l

+ U>

0>0 «i -a

«0 1

V -t w> • t •H •a

c- l-l 1 .

IO rM

1 ,

•H 0

o to

o CM •-I

o l-l 03 •H 0

o IO CM

O «l IO

o s ^1 i1

O 3

O

10

s I' •0 *l IO

0»

I i- •* •0 CM

O to a «

•H O

8 l-l

O IO t-

o

|1 Is o l-l CM

o iH CO

1

o II a H

o to

7 o

%

»< S

o C-

O 00 *l §

Is o 00 CM

8 CM 1

ll 8 o

0»

i! + * «0 CM + • •f •f

M> f t V y> -is 1? to * # • • • "I + • • • "? T

1- B 0 o •Ml CO

o M IO

o CM IO

3 si 8 •H

O CO IO

o 2 h i-

o Ö §

t- o us

Ji- *• *• IO

«0

s «

IO •* IO

t-

•H l-l

ll o o o I- o O o s o ft to fl B g IO c- o 10 f CO «4 C-

•H § a K

CB to v> pa rH •

I- 2* C 1 •0 «4 O • 1 1 •a a ü 8

a o 8

a a s a t-< ^

li Tt

0 •P a It

o CD

o o •rl

3 • •SI n g

CM 1 1 «•

MJl

1 V 4>

O l-l «4 CM +

• • 1 Ml« 7 1

to

<*4 1 -v -r tt • J! f +

m H "O 0 H "O 0 j •« •

• 0

^-^ 0 4»

0 0

•»«•

0 g. O o o J» o o 8 ^1

O o O U

a 3' io CO

» CM IO

IO m O u m c-

i IO c- CO

0

1 CM

3 5

•

I OS

fl

8 to

"?

8 CM

1 8 o> 1

r-t 0 •o

I 0 hi

1 If

8 OS i!

a A

o IO

o l-l CM

7

o 8 u • IO 0J

8 03

•H O a a

o CO o

! c.

J fl

o l-l

n s CM M

•f

o m CM CO

• il 8 IO

HV

8 8 •

o CO CO

M u

I «H 1 1 li

o r* to

• §

^-^ •£ H|« ^-s 1 Hh» ,—N

w u iH IO Ml M> co =1 4> %c 8 0» *rl • « CM CM •H «4 CB 0 1 "? 1 • Ä W •4 T) + + * tö 1 l + +

1* O CM

O r-l rH

8 t- »H 0 !

o s

o IO e-t

I» o 3

o M «0

to

s

•H CM iH

«0

3 M

J- f» 10

04

!«

o l-l

O 2

o 0> li o

CM O li o

CO IO

o to 2

as ¥ * % 33 •f

•rl O a fl + +

J5

o to 8 o

l-l o h 8 IO

o c~ IO

a IH 1 a 00

o 8

OS

• 0» li •M"

+ + ll •H

• + ,-__,

»• +> • • +> • •

1 n o I u

CO • m •

o 0

fc -1 CO • •<

H) •H o T-t M o M

§>/ •4» o +> O

o

B-3 • / HI /

+> •8 •s"i • / * *> O

Ä/ 1* $ US C ho M & / CJ3

• o5 fi -H o 3 IO M O «4

»H

o »•

/ • kc °c <OCM 1 8 °c °c / S °C °C

/ S? 5 O 0 -i 0

• O CO

IO •

/ ** / 4> 4>

/ S 8 O 0

J3 P.rH

IO 0

• Oca

4» •

loi / 2s / •H ° O « to •

tHr4 »0<-l lOM •Hr-I •0 rH IO M / C c rtr-

•H l-l rH / Bb O •-I •H l-l

«H -H fc


TABLE IV.« 3PANWI3E AND CHORDWISB STRESS CHANGES AT ROSETTE STRAIN- GAGE LOCATIONS IN WING OUTER PANEL OF THE G-k6 AIRPLANE RESULTING

FROM OPERATION OF THE THERMAL ICE-PREVENTION SYSTEM

High-temperature delta rosette« at station Itf

Flight ^^-^^ eonditlon ^""""^-^^

R05 RG6 RO7 RG8 Span- rise

Chord- wise

Span- wise

Chord- wise

Span- wise

Chord- Span- wise wise

Chord- wise

110 mph eorreet I.A.S, level flight ••639O •855O -366O 43I4IO «418O •5470 -3100 •U750

135 mph oorreot I.A.S* level flight •5890 •898O -3250 •3190 -i&OO +4580 -3720 4J4J+6O

155 to 165 mph correct I.A.S. 2g bank •57IO +8750 -3820 •3250 -59OO •4950 -14.020 436OO

High-temperature delta rosettes at station 137

^"-"~\_^ gsg«

eonditlon ^"~"~-~-^^^

RO36 R037 RG38 R039 Span- slse

Chord- wise

Span- vise

Chord- wise

Span- wise

Chord- Span- wise vise

Chord- wise

110 mph correct I.A.S. level flight -357O +5^00 -69UO +3230 -50ltf •3180

135 mPh correct I.A.S« level flight

•I7OO •3310 -2210 +5300 -617O 42990 -lj.810 42260

155 t° 165 mph correct I.A.S. 2g bank •2630 +3665 -I65O +788O -788O 4318O -33U0 +47OO

Low-temperature 1|5° rosettes at stations I4.7 and 137

Flight ^^^>-»^. condition ^"^-^^^

RG4 RG32 RGi|l R060 Span- Chord- Span- Chord- Span- Chord- Span- Chord-

110 mph correct I.A.S. level flight •1630 4830 -570 4310 42080 •120 -230 4790

135 BP° oorreot I.A.S. level flight •1280 •16IO -1*20 •390 •I960 --30 4260 489O

Note; 4 denotes tension stress, pounds per square inch. - denotes compression stress, pounds per square inch.

Data corrected for Polsson's ratio effect.


TABT.H! T.- TBJPEBATUBB CHANGES DT WUTG OUTEH PAHEL 0? THE C-46 AIBPLAHE SBSULETDIO JBCM OPERATION 0? THERMAL ICE-PHEVENTION SYSTEM. PAST 1 - STATION 47.

Tamporatura °F • "^Flight

^xoondi ti on«

Tharno- ^s. oouplaa ^«^

110 nph I.A.S. lOTOl flight

156 nph I.A.S. 156 to 166 nph X.A.8. laral flight 2g bank

Boat off

Boat oa

Boat off

Boat on

Boat off

Beat on

Tl-1 48 46 46 47 46 60

Tl-2 48 46 45 47 46 45

T2-1 46 49 46 50 46 48

T2-2 46 49 47 50 47 62

T5-1 47 68 47 54 47 56 T5-2 46 68 46 64 46 86 T4-1 - 47 88 47 64 47 87 14-2 46 64 47 65 46 67 n 61 129 51 120 48 152 T7 66 161 65 167 47 168

18 84 124 60 124 60 152

n 49 74 60 76 49 75

T10-1 80 66 50 67 50 76 T10-2 60 65 50 66 60 70

Tll-1 61 64 61 66 81 77 Tll-2 60 66 49 66 61 65 T12-1 49 65 49 66 61 60 T12-2 48 62 48 65 80 67

T1S-1 47 49 48 61 64 86

T15-2 47 49 47 49 84 64

T14 46 48 46 48 47 61 T15 47 60 46 61 48 84 T16 47 60 48 62 48 66 tn 49 65 48 64 48 58 T18 46 56 48 66 48 61 T19 49 67 49 58 48 65 120 49 68 49 67 47 75 T21 64 94 65 97 50 106 T22 61 69 61 70 50 80 IBS 69 75 54 77 51 88 XI« 68 75 65 76 82 86 T26 62 65 52 70 62 80 T26 82 59 65 65 88 71 »27 49 65 50 56 62 60 ft* 48 65 49 56 68 69 f» 48 68 49 66 64 60 ISO 47 80 49 55 64 68

Wl 47 49 48 82 64 86

lot« i 1 danotos tharmoooupl« on laaor 2 danotc« tharnooeupl« an outar

•urfaoe of akUu aurfaoo of akin.


TABLE V.- TEMPERATURE CHANGES IN WING CUTER PANEL OF THE C-46 AIR- PLANE RESULTING FROM OPERATION OF THERMAL ICE-PREVENTION

SYSTEM, PART 2 - STATION 137.

Flight condition

Tharaw- ooupla«

T32-1 T32-2 ToTT "TOTT TS4-1

T3T=T TBT "T5T T3T T40" TTCT THIS"

T42-2 T4T=r T4S-2 T44-1

T46 T47 TIT T*T TBT TST "TBT TTT "TBT TBT TBT TBT TBT

TBTET TeTT

Taaporatnro °F

110 mph I.A.S, lml flight

Haat off

64 55

-5T TT TT "BT TT TT TT 47 TT

TT 48 TT 46

TT TT 46

bat on

66 66 TT "6T "7T TIT

TIT W TUT

89 TT T4 TT 60

TT 64

-EÖ- "BT "4T

135 »ph I»A.S. lava! flight

Haat off

50 To- To" TT TT TT TT TT TT TT IT TT TT 48 TT 47 TT TT TT

46 46

TT IT TT TT 50

"BT TT TT TT TT TT 47 TT TT

63 TT •6T TT

"ST "ST 86 77 TT "BT TT "BT

4« 46 TT TT TT TT TT 60 TT TT TT TT

TT •w TT

TT 48 TT TT

Haat

66 65

TT TT

"TT "TT W

TUT IT TT TT TT TT TT TT IT TT 48

64 55 TT TT 95

TUT 86

TT TT 68 TT TT TT 61

TT •TT

155 to 166 »ph IJL.S 2g bank

Boat off

40 41 TT TT TT

[6 TT TT TT TT 48 TT TT TT TT TT TT TT

47 47 TT TT TT TT TT TT "BT TT "BT "BT "BT TT TT 46

Haat on

46 46 TT TT TT "BT

TST "rar TOT "TT TT TT TT TT TT TT TT TT TT

6T TT TT TT

"IoT "IDT TT TT TT TT TT TT TT TT TT TT

1 Danotai tharnooouplo on innor ourfaoa of «kin,

2 Danotas tharnooouplo on outar ourfaoa of skin*

26 KACA ARR No. 5G30

TABLE V.- TEMPERATURE CHANGES IN WING OUTER PANEL OP C-I4.6 AIRPLANE RESULTING FROM OPERATION OP THERMAL ICE-

PREVENTION SYSTEM. PART 3 - STATION I59

SZ/

A/9

*#*#*

,4 /a Thermocouple designation A - air S - akin M - structure

Temperature °p ^\Plight

^-vcondi tions

Thermo- ^v oouples ^\^

110 mph I.A.S. levelflight

135 mph I.A.S. level flight

155 to I65 mph I.A.S. 2g bank

Heat off

Heat on

Heat off

Heat on

Heat off

Heat on

321 Ä 151 62 lhl ?2 157 SI9 59 I52 5« 11+) 47 1$B S20 55 147 54 124 48 l6l S23 * 150 52 lU8 49 164 S25 53 iks 51 150 50 160 A19 52 125 5U 134 50 151 AI3 58 256 56 247 50 269 Al8 51 131 50 134 50 11*2 M5 59 216 56 214 52 21+4

NACA ARR No. 5020 37

TABLE V.- TEMPERATORE CHANGES IN WING OUTER PANEL OF C-Ij.6 AIRPLANE RESULTING FROM OPERATION OF THERMAL ICE-

PREVENTION SYSTEM. PART I* - STATION 380

ASZ

J0Xa*&S?0

-S+/ LJ'^3^/4S3

Thermocouple designation

A - air S - akin

Temperature °F ^.Fligbt

^"vpondi tions

Thermo- ^v couples ^\



155 to 165 mph I.A.S. 2g bank

Heat off

Heat on

Heat off

Heat on

Heat off

Heat on

Slf2 1*9 112 5k 1U5 k9 157 SlfO 51 137 57 i6i 50 176 S39 1*7 169 kl 163 1*7 180 Sl|l 1*7 I6l k9 157 kB 167 sltf 1*7 155 1*6 11+6 kB 138

A32 kl ll*9 kl 152 1*8 160

A27 k& 278 50 26k 1*7 286

A53 kl 180 1*6 nk 1*8 179

NACA ARR No. 5G20 Fig. 1

so *

I

a <D -P 00 >» ra

a o •H -K> Ö CD

M fX I

CD o

CO 8 H <a X}

a)

> -s /

•Ö CD P. P. •H

D CD

<D

s P.

•H cd

CO

I o CD XI £-<

N" 1

CD

ä •H

NACA Fig. 2

NACA ARR No. 5030 Fig. 3

IATI0IAL AD7I80RT C0HMITTI1 FOR A1R0IAUTICS

3

iJACA ARR No. 5080 Fig. 4

i

NACA ARR No. 5G20 Fig. 5

<

to H a o

«rl

4» CO

i-H « s M

O

2

$

8 o e © *x •»

s to

C8 M 4»

«H o C 0

«5 p. o u °ri I

«

&

NACA ARR No. 5030

I

S)

As-A s CP3^;

s!1

I §^

I *

*fc

*W^ $

I ^l*M .•XN»I

1 V i

Fig. 6

63 M

*« "51 S

aß S

NAOA ARR No. 5G20 Fig. 7

Figure 7.- Typical strain-gage and thermocouple installation In left-wing outer panel, C-46 airplane.

NACA ARH No. 5G20 Fig. 8

Ct» a u CO

-u

bQ Ö

•H tS

CD

+3

u o

a) o CD

a M •H cd •

Ö Ö •H cd

rH Ö ft o n •H H V CtJ CO

•p CO

CD I a o

^ •>

CO o

•H +J 08

«H bO O «H

•P C CO o «

•H > •P Ö CO -H .H rH CO a) co +3 a> CO h fl-p H CO

I

00

CD

3>

NACA ARB No. 5020 Fig. 9

3 4

oooo 0 occo \ 5!oooo

^ ^ocoo ooos cooo oooo ccoo

N »IOOCC

^ Moooo j Noooo ^ '«oooo

oooo OO CO

oooo oooo oooo oooo. cooo cooo

^ Socoo X \cooo ji '"lOOOO

s Noooo

oooo ccoo oooo OQOO oooc ocoo

OOOO

OOOO

N \3:oo oooc ocoo S N \ •* * * s s 1

X

X

*s«

3

83 £| «o

Is M

5

c e > to

4»

o u

•» • u

§•1

e^ss

o __

4» •

gg «H O O

St •r« M « 60 U O O iH OH

M O OB

I« boo C a «i bO * C

.5 0

4» O

«2 O bO

a« ^ a» 0* -HO" 4» O <4 • a H o o o OHH

I

0> o

I Ü1

VAC A ARR »o. 5030 rig«. 10,13

3

I. •3

•o o

H • •r* a • »4 » B a o OVl

9~. \

COCA

3

• o u Vl

o

3 o

2 ' 3 i

• 3 h 4» •> O • 3 u

•H 4» M

dw

3 5 Vl • e • Od o 3 4*

•H rH 0 m o MB • • • 35 B • 3

o Sn? o a •H f« O CO c h 3g 13 M 3 ^3 «4

5. rt^ § • 0> O * u *> Pi _ •>* B %4 o

*4 • 8 o

1

a h a • 9

43 •1 • 3 o o ' 00£<fc •^ \

1

3 a • u

i . as

• hr-4

8 i «» • s • B u •

4» H

\ \ •s 8 "8

e p. «4

N • • f!

\ 38 ° ^J " °s%

t 8 t \ 0M t-* -H 4» \ •* * * •

t- o \ A to O >» 5 | 3 | 0

\ a d o •

o ho o «4

4» x> 1 •> fj a n •a

U to «

o •*«. & e

h. -rt

IATI0IAL ADVIB0RT COMHITTII TOR AXROVAUTI08

\ \>

\ V \ vs

{s ^

V*N.

a $

•^—1 >

• M 5»

, is.

2 2.

; ; . \l\

1, 4» 1 '2S •-Of

Hi-

o

• sS

^ s

• "v ">

^ \

44 4»

3 <s ̂

\S

o •»5

i to

44

1 SP M

a •H •

to o n

bo a? H3 Vl • O M

d U

s« -° 5 9 . -it* O MO

»o >• O-f- M B O »4» 4j 4> «

U Vl 4»

3°S 4 4» 4» HOI

• d HVl O «Vl O o o

H-4 4J

s f

11 TM '«««a*«

NACA APR No. 5G20 Fig. 11

t

i 00 CO <D M +3 03

<H O

>» o

u 0 • O <D

cd d

a) CD J4 O V> o

M 0) p,

CO CD

CO +3 +* I <D +3 •0,3 o -H -PrH

<M •Ö <D -Ö

SS ß+3 O Ö -H CD +> H a) P, .H -H

cd er +» CD 0) fl+» •H Ö

CD +J s CO <D CD M

+=> 2 CO

M CD o a cd CD

o M §

cd .a

CD u s,

•H

J IAUA AMH •0. biiS 0 rig. 12a, b

4 rH 10 2 >•

3 < »1 j •I /

/ a o

2& • s /

/ 0 o

Si /

V»

9 / u 3

"> P. •

8 Sa r|.rl

co m x +

• 9 /

»4 0 1

' V

«0 1 rH «

o - 5

CO § 1 s rtM s

O A 11 2"" S o 5 .d ! „

gt> a

/ / 1* t- } s - ^

{,' •

j r to

l{ c o O -H

• •A 7 V O M

o *> • co co 4

0 fl 8 •

S i / /T3

1 -J*

7" /co J M VI

° M.

—. V

^ • in / c? •*•

/ ̂ V

V s ^+- 3 » ra O • " <0 * M O CM »i

TM '•»«•tie' •«oj« T»n <t)So«T(o IIUU

O O O O O O QOOOOO O CO (O •* CO O 0D to •* CO rH H i0 'aSovno •ir\v><i«dm*£ J0 '»9u»qo ejTt^vjedmsj

lATIOIAL ADVIS0RT OOIUIITTXX TOR AJROllAUTICe

9 •

co f o

c *4

n o •

m H

< • -

• 2

/ t

• 9

• o -9 *» «

f>

"I • •» • • 0 h-H 4» M CO CO

X +

• / 3

> / H

3

• it f

/ ri • / B 9 - -2

«4

J» /, £ 1 1

J3 3 CO £>

o /, ;3" «rH U

O .grH ^

1 CO g5 s

1 J ^ i tst- 5

a a o o • •s .-'

.- JJ

'f CO

\ r +

A I '

?co sf i i I L •H

CO -j o

is 4

2 « y in CO

rt rH

•

s ^+' "—"«=;

coX v«- f

^- "-*> v*-

8 o CD

>3wr

o co

CO

o

191%

O CO

o

B

o

CO * 1

0

8

'•i

o 00

lu»qs

o CO

m

> ••(

O •H»

»«a

o CO

o

o

t 1

4C '»8u»t(0 »jrn»j»dm«i i0 '»9tr»r(0 9iTUv.x»dac»j

HACA ARR Mo. 5020 Tig. 12c,d

«0

s 10

0

•

a X

-i ft «

£* 0 »

c- 111 i /

0 •I 1 1 / •

1 0 • 0 m~ 0

/ u 3 /

<H u 9

/ • / /

1 . 8 •

c t- V

a •rl M

1 /

5 CD »CD

X-*-Q

1 * /

f 0

! r- x 7 i

t <0 / / 0

\ /? \

Y

i n 0 •a in \ /

>

\ ,p 1 0 1 5,10 \

X \ ^ r r ^

r4 »lO a" 7 0

J2- J s« > /

0

H / <- s

r

' X- 0> 10

at 10 / / 00

10

c a 1 3 ' C y T •? ' « < 3 ' < 1

» ?

I

*> 0 O0 o>-t H

T8* '»Swrqo t«»Ji8

800000 co <o •* c»

i0 '«9u«qo »2Ti}».i«dla»i

t»X '»Stntqo asejis

00000 CO «3 •* o»

•ATIOIAL ADVISORY COUCITTIX FOR AIROHAUTICB

8 1-1

t0 'sÄtrtqo »jTt}«j»da*x

d o o

H to

3 +» 0 0 h U 3 0

si- M 0 0) +i O. S U O <u

Si < O X

to )

; / /

./ 0 3

1

/ 0 o_

/ g / •r

a CB >

f M O

1 0

h 01

Str

« S

kit

8tr

i

A 9 ll K 1 O

X + 0

It %

\] \i

•

r

r/X •

s

•

.1 7 to-

ol ca S

+1 CO« 1

i-t

w 0 •9

/ S) //

1 t! ^ / ?/o>

I c"-„ -J 03

1 1/

0/

1

/ in /

„ + CO

x-^

-^f- —X to

3 <

<• ««St

D < X) 1

( T8TJ0

3 1 0

88«

3 C * <

H8 O

0

t 1

* r c c

> < 3

* •eSt

3 < 0

( rsqo

D C O •<

a 880.

3 < * C

<

> a

O

c a 1 T

o to

o in

m to

o

1

m in

h

:^ ' o

•i? o*a >S -

o

i0 'aSuaqo sjn^jsdmei i0 'aSvaqo 9jrn»jedm»j

HACA ARR No. 5020 rig. 18«.t -r • 3 8 lO

o H.S

M h 3 • 1

9 •-

Si «•

• <

i i / 1 o

* / 1 1

co

• a o a-rt 1 U-t n /

r

s 1 1 €>

o «OODCO /

X+O j 3 / 1 1 3

m

r • / 1 1 e s

\ 1 1 i 1

o •5

• •2 \

+

/ 1 °° ') o I - in

\ i1 \

1 /

+ 1

i <! / /

o /

o

\ / /

/ / j W

TO IOXC

r '+ XO 5 ,

in ^* / O

/ <£ H

-^ -X 0) 10

to Ns c V] i c

1 T I c ? 0 0

1 ( » o TO

1 * f «r

U

& in <o

33 o a

o

I«* *98tr»no ssajq.8 T«t *»8tmt(0 lltits

o co

o CO

o TO j '93u»qo »jnt»j9dmex

8 o CD

O CO

o o TO

3 a •» a 8

TO IATI0IAI, ADTI80RT COHHITTSI

FOR AEROHAUTICS

T8JI '«Santo »is-ns f«3t *e3xr»qo sesame

8 o CD

o CO

o o o •* TO o CO

o CO

j *a8avt{c ajTuvjsdmai J 'SStTBHO 9inj»j8dw»i

NACA ARR No. 5G20 Fig. 13g,b

\an '»3tr»qo easing \B^ 'sStrsqo saa^is

NATIONAL ADVISORY COMMITTEE TOR AERONAUTICS

CM O

ISTi 'a9u*qo assj^g

* N O W

fB^ 'aStrstjo sssj^s

r TITLE: An Investigation of a Thermal Ice Prevention System for a C-46 Cargo Airplane. VII - Effect of the Thermal System on the Wing Structure Stresses as Established in Flight AUTHOR(S): Jones, Alun R.; Schlaff, B. A. ORIGINATING AGENCY: National Advisory Committee for Aeronautics, Washington, D. C. PUBLISHED BY: (Same)

OTD- 8200

_LHr>sl xusC, SJ "I KRPßfsO

PU2USJCX3

Sept '45 Unclass. U.S. PACES

42 tUUnSATIOMS photos, graphs

ABSTRACT:

The thermal ice prevention system is investigated to determine change in stress at various locations in wing outer panel caused by operation of thermal system. Wire resistance strain gages and thermocouples were installed at numerous points of wing. Change in stress and temperature are obtained for two speeds in level flight and one speed in 2g bank. Curves show variation of stress and temperature for two wing stations. In order to evaluate magnitude and seriousness of stress changes, comparison is made with allowable stresses for the unseated wing.

DISTRIBUTION: Request copies of this report only from Originating Agency DIVISION: Stress Analysis and Structures (7) SECTION: Structural Theory and Analysis

(Methods (2)

ATI SHEET NO.: R-7-2-13

SUBJECT HEADINGS: Wings, Heated - Stresses (99264); Ice formation - Prevention (50201); C-46 (19530)

' Documents Division, Intolligortco Deportment Air Materiel Command

Aid TECHNICAL INDEX Wright-Pattorson Air Forco Bcao Dayton, Ohio

-i

4

UHCIASüIFED PER AUTHORITY; INDEX Ül- KüCA TI-XHKICAl PUbLlCATIGIiö DATED 31 DECEMBER 194?»