— RM L58B19 NACA RESEARCH MEMORANDUM EFFECT OF COMPRESSIBILITY ON THE HOVERING PERFORMANCE OF TWO 10-FOOT -DIAMETER HELICOPTER ROTORS TESTED IN THE LANGLEY FULL-SCALE TUNNEL By Joseph W. Jewel, Jr., and Robert D. Harrington Langley Aeronautical Laboratory Langley Field, Va. UNIVERSrTY OF FLORIDA DOCUMENTS DEPARTMENT 1 20 MARSTON SCIENCE LIBRARY RO. BOX 11 7011 GAINESVILLE, FL 32611-7011 USA NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON April 21, 1958 Declassified May 29, 1959 NACA Rl-1 L^8B19 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEARCH MEMORANDUM EFFECT OF COMPRESSIBILITY ON THE HOVERING PERFORMANCE OF TWO 10 -FOOT -DIAMETER HELICOPTER ROTORS TESTED IN THE LANGLEY FULL-SCALE TUNNEL By Joseph W. Jewel, Jr., and Robert D. Harrington SUMMARY An investigation of the effects of compressibility on the hover ing- perfomance characteristics of two 10-foot-diameter helicopter rotors having solidities of approximately 10 percent has been conducted in the Langley full-scale tunnel. One rotor, having NACA 0012 airfoil sections, a plan- form taper ratio of 3/l> and. -8° of twist, was tested to a tip Mach number of 0.95 an d. a disk loading of 16 pounds per square foot. The other rotor had NACA 64-series airfoil sections, tapering to a 6-percent-thick tip, a plan-form taper ratio of 3/l, -l6° of twist, and was tested to a tip Mach number of 1.0 and a disk loading of 20 pounds per square foot. As the tip Mach number was increased to 0.86, the NACA 0012 rotor encountered compressibility losses at progressively lower thrust coef- ficients. At tip Mach numbers of 0.91 and above, compressibility losses were evident at the lowest values of thrust obtained in the investigation. No effects of compressibility were evident at any thrust coefficient tested on the NACA 6k- series rotor up to tip Mach numbers of 0.91- At tip Mach numbers of about 0.92 and above, compressibility losses were evident at the lowest values of thrust measured. The measured power requirements of both rotors at the lowest tip Mach numbers are sig- nificantly more than predicted by usually reliable theory. This effect is believed to be due to the unusually large ratio of hub diameter to rotor diameter. The absolute values of the measured data are therefore not believed to be representative of full-scale rotors; however, the effects of Mach number on either of the rotors or comparisons between rotors are believed to be valid. The maximum sound level measured in the tunnel test chamber at a point 21.5 feet and lU0° from the vertical shaft on the NACA 6^-series rotor was 122 decibels at a tip Mach number of 1.0. NACA RM L58B19 INTRODUCTION One method of increasing the forward speed and lifting capabilities of the helicopter and thus its utility is to operate the rotor at high tip speeds. Accordingly a general research program is currently under way at the Langley full-scale tunnel and the Langley helicopter test tower to determine the effect of compressibility on the hovering and forward-flight performance characteristics of helicopter rotors. As one part of this program two rotors 10 feet in diameter have been designed to study the forward- flight performance characteristics of rotors operating at high tip speeds. Hovering tests of these two rotors have been made as a preliminary step to aid in the interpretation of the forward flight results as well as in evaluation of the hovering performance of rotors designed for high forward speeds. This report pre- sents the results obtained from these hovering tests. One set of rotor blades tested had an NACA 0012 airfoil section, a plan-form taper ratio of 3/lj and -8° of twist. This rotor was tested to a tip Mach number of 0.95 ajn -& a disk loading of l6 pounds per square foot. The second set of blades had symmetrical NACA 64-series airfoil sections, tapering in thickness ratio from a 6*4-2-015 section at the cen- ter of rotation to a 6^-006 section at the tip. These blades had a plan- form taper ratio of 3/l and -l6° of twist. They were tested to a tip Mach number of 1.0 and disk loadings to 20 pounds per square foot. For this rotor, sound-level measurements were taken at each test condition. SYMBOLS a slope of section lift coefficient against section angle of attack (radian measure) b number of blades per rotor c blade section chord, ft / cr^dr equivalent blade chord (on thrust basis), — — , ft I, R r^dr NACA RM L58B19 c^ section profile-drag coefficient Cj section lift coefficient 6Cm cj mean lift coefficient, — - Cq torque coefficient, jtR 2 p(aE) 2 R Cq ^ induced-drag torque coefficient, Cq q profile-drag torque coefficient, rp Cm thrust coefficient, Qi jtR p(nR)^R Op *R 2 p(fiR) 2 R rtR 2 p(^R) ; I mass moment of inertia of blade about flapping hinge, slug -ft M rotor figure of merit, O.'JO'J- 2 C T 5/2 M^ blade-tip Mach number (ratio of blade-tip speed to speed of sound) Q rotor- shaft torque, lb -ft Qj_ rotor induced-drag torque, lb-ft Qq rotor profile-drag torque, lb-ft r radial distance to blade element, ft R blade radius , ft T rotor thrust, lb v induced inflow velocity at rotor, ft/sec k NACA RM L58B19 x ratio of blade -element radius to rotor-blade radius, r/R 0 at any Mach number to the incompressible profile -drag torque coefficient as a function of rotor mean-lift coef- ficient c^. Such a plot is shown in figure 6. The profile-drag torque coefficients were determined by subtracting a calculated induced-drag torque coefficient Cq ^ from the corresponding measured total torque coefficient Cq. It is apparent that as the tip Mach number increases the rotor encounters compressibility losses at progressively lower mean rotor lift coefficients. At tip Mach numbers of 0-91 and greater, the rotor is operating well above the drag rise even near zero mean lift coefficient. These same profile-drag torque-coefficient ratios are shown as a function of the calculated rotor-blade-tip angles of attack in figure 7. From this plot it is possible to estimate the combinations of tip Mach number and angle of attack at which drag divergence occurs on the rotor. A comparison of the drag -divergence Mach numbers estimated from figure 7 with those indicated by two-dimensional airfoil data is shown in fig- ure 8. In general the rotor drag-divergence Mach numbers are about 0.10 higher than those shown by the two-dimensional airfoil data. This value compares with values of about 0.02 to 0.06 which have been measured in previous tests (refs. 2, 4, and 5) • Comparison with theory . - The experimental and calculated hovering - performance curves for this rotor at tip Mach numbers of 0.45, 0.73, and 0.8l are compared in figure 9- The calculated results are based on a strip analysis (see ref. 6) in which section lift and profile-drag coefficients were varied with Mach number along the blade. The varia- tion of lift and profile-drag coefficient with Mach number used for these calculations is shown in figure 10. Two different tip loss factors were applied to the calculated thrust and induced torque. One of these, B = 1 - -^ , is based on a triangular inflow distribution, whereas b 1 i[£t7 the other, B = 1 - — — V § , is based on a uniform inflow distri- bution. The development of these equations is shown in the appendix. NACA KM L58B19 Because of the twist and taper this rotor probably has an inflow distri- bution somewhere between triangular and uniform. In addition, at tip Mach numbers of 0-73 and 0.8l a tip Mach number relief was included in the calculations. This relief is characterized by an effective reduc- tion in tip Mach number due to three-dimensional flow at the blade tip. It was assumed that the effective tip Mach number was 0.05 less than that calculated from test conditions and that this relief varied linearly to zero at the 3 A blade radius. The more usual value of 0.05 was chosen for these calculations rather than the value of 0.10 indicated in fig- ure 8. Inspection of figure 9 shows the relative effect of each of these correction factors on the performance calculations. The theoretical performance calculations made for this rotor at a tip Mach number of 0.1+5 predict an average of 6 to 9 percent more thrust for a given amount of power than was actually measured depending on the type of inflow distributions assumed in the calculations. (See fig. 9(a).) This agreement is considered only fair when compared with the close agree- ment generally obtained at low blade-tip Mach numbers on larger diameter rotors. (See, for example, refs. 2, 3> ^> and 7-) An explanation of why the measured rotor performance is significantly less efficient than the usually accurate predicted results has not been precisely defined. There may be several factors present in the test setup, however, which have a detrimental effect on the measured rotor performance. It is believed that the primary reason for the discrepancy is a flow disturbance caused by the unusually large ratio of hub diameter to rotor diameter, which in these tests was about 0.2. For this reason it is believed that the measured results are repre- sentative only of rotors having similar ratios of hub diameter to rotor diameter. The calculated results are believed to be more nearly repre- sentative of rotors having smaller ratios of hub diameter to rotor diam- eter and similar hub configurations. However, it is felt that the effects of Mach number on either of the rotors or comparisons between the rotors are valid. Somewhat better agreement between experimental and calculated per- formance is indicated at tip Mach numbers of 0-73 and 0.8l. (See figs. 9(b) and 9(c).) It is not known whether this better agreement is fortuitous or whether the effect of the flow disturbances due to the hub are mini- mized because of the presence of supercritical flow existing at these higher Mach numbers. Symbols were used to identify the different methods of calculating the theoretical performance . Power Requirements of NACA 64-Series Rotor Performance measurements .- The measured variation of thrust coef- ficient Cx with torque coefficient Cq over a range of tip Mach numbers NACA RM L58B19 9 from O.lj-5 to 1.0 for the rotor having the NACA 6l+-series airfoil sections and -l6° of twist is shown in figure 11. For this rotor no compressibility losses are evident up to and including a tip Mach number of 0.91. At tip Mach numbers of 0.95 and 1-0 compressibility losses were measured over the entire range of thrust coefficients covered in the tests. However, the characteristic spreading of the performance curves is not evident even at the maximum thrust coefficient obtained (dp = 0.0068) at a tip Mach number of 1.0. This point corresponds to a disk loading of 20 pounds per square foot. The reduced compressibility effects measured on this rotor, compared to the NACA 0012 rotor, are attributed to the differences in thickness ratio and blade twist. Profile-drag power ratio .- The profile-drag torque ratios for the NACA 0*4- -series rotor as a function of rotor mean lift coefficient and calculated tip angle of attack are shown in figures 12 and 13, respec- tively. The profile-drag torque ratios remain constant with tip Mach number and rotor mean lift coefficient or tip angle of attack to a tip Mach number of 0-91- At tip Mach numbers of 0.95 and. 1.0 there is an increase in profile-drag torque ratio. However, the ratio remains rel- atively constant over the mean lift coefficient and tip angle -of -attack range which indicates that the profile-drag torque is independent of tip angle of attack at a given Mach number over the range of the tests . The combinations of tip Mach number and angle of attack at which drag divergence occurs on the 64-series rotor have been estimated from the data of figure 13. Figure lU presents a comparison of the experi- mental drag-divergence Mach numbers with those indicated by two-dimensional airfoil data. The experimental drag-divergence Mach numbers range from about 0.08 to 0.11 higher than two-dimensional results over the range of tip angles of attack of the tests . Comparison with theory . - The experimental and calculated hovering - performance curves of the NACA 64-series rotor, operating at tip Mach numbers of 0.i+5> 0-95> and 1.0, are compared in figure 15. In general, the agreement between calculation and experiment at a tip Mach number of 0.^5 is similar to that obtained for the NACA-0012 rotor. As indi- cated in the discussion of the NACA 0012 rotor, the better agreement at high Mach numbers may be fortuitous. The calculating procedure and cor- rection factors used in the analysis of the 6k- series rotor are the same as those used for the 0012 rotor. At tip Mach numbers of 0.95 and- 1-0 the outboard blade sections were operating above the limits of the available two-dimensional airfoil data. Therefore, it was necessary to extrapolate the airfoil data to a Mach number of 1.0. This was accomplished by using the results obtained on similar airfoil sections, tested up to a Mach number of 1.0 in a differ- ent wind tunnel, as a guide in the extrapolation. The two-dimensional 10 NACA RM L58B19 lift and profile-drag data used in the performance calculations for the 6i+-series rotor at tip Mach numbers of 0-95 and 1.0 are shown in fig- ures 16 and 17. Comparison of Compressibility Effects on the NACA 0012 and NACA 6U-Series Rotors Figure of merit . - One method to determine the relative character- istics of two rotors of the same solidity is to compare the variation of their figures of merit M with the ratio of thrust coefficient to solidity Op/a at equal tip Mach numbers. Plots of figure of merit as a function of Cp/a for the NACA 0012 and 6k- series rotors are shown in figures 18 and 19, respectively. Comparison of figures 18 and 19 indicates that at Op/a less than 0.0^ and at tip Mach numbers up to 0.8l there are only minor differences in figure of merit between the two rotors. At Op/a greater than 0.05, the Gk- series rotor is superior at tip Mach numbers of 0.8l and above. At a tip Mach number of 0.86 the 64-series rotor becomes more efficient than the 0012 rotor for all values of Op/a greater than 0.02. Since the solidity of the two rotors is almost identical, these differences in efficiency must be attributed to the use of the thinner low-drag airfoil sections and to the increased twist. At the lower values of tip Mach number the 0012 rotor would be expected to have somewhat higher figures of merit because of the better lift-drag ratios of the blade sections at higher lift coefficients. How- ever, at the higher tip Mach numbers the thinner tip sections and lower tip angles of attack resulting from the higher twist tend to delay the compressibility effects. Thrust-horsepower ratio .- A dimensional way in which rotors of the same solidity can be compared is on the basis of the variation of thrust- horsepower ratio with disk loading. Plots of this type for the NACA 0012 and 6U-series rotors are presented in figures 20 and 21, respectively. In general, as the tip Mach number increases, the maximum thrust-horsepower ratio is reduced. At the same time the variation of thrust-horsepower ratio with disk loading is reduced and the optimum value for a given tip Mach number occurs at a higher disk loading as the tip Mach number is increased. The improvement in hovering efficiency at high blade tip Mach numbers brought about by the combined benefits of reduced airfoil thick- ness ratio and increased blade twist is indicated by comparing figures 20 and 21. NACA RM L58B19 11 Sound Measurements It is well known that the sound emitted by rotors and propellers increases in intensity with increases in tip Mach number (ref. 8). As a matter of interest, sound-level measurements were made at each test condition for the NACA 64-series rotor. The results of these measure- ments are shown in figure 22 as the variation of sound pressure in deci- bels with tip Mach number for several values of thrust coefficient. At a tip Mach number of 0.^5 the sound pressure varied from 82 to 9k deci- bels for the test range of thrust coefficient. At a tip Mach number of 1.0 the maximum sound pressure was about 122 decibels. As a matter of interest, some calculations were made of the "Gutin" or discrete frequency noise from this rotor by means of the method out- lined in reference 8. The results of these calculations are shown by the dashed line of figure 22 which is a summation of the first five-blade passage frequencies for a value of Op of 0.006. It can be seen that the general trends of the calculated and measured data are similar and that the sound pressure levels agree within about k decibels. It should be pointed out, however, that the method of reference 8 does not predict random noise levels. Hence, for conditions where the random noise is important (full-scale rotors having large blade areas) this method is probably not adequate. CONCLUDING REMARKS An investigation of the effects of compressibility on the hovering- performance characteristics of two 10-foot-diameter helicopter rotors having solidities of about 0.10 has been conducted in the Langley full- scale tunnel. One rotor, having NACA 0012 airfoil sections, a plan-form taper ratio of 3/l, and -8° of twist, was tested to a tip Mach number of 0.95 and a disk loading of l6 pounds per square foot. The other rotor had NACA 64- series airfoil sections, a plan-form taper ratio of 3/1, -l6° of twist, and was tested to a tip Mach number of 1.0 and a disk loading of 20 pounds per square foot. As a result of this investigation the fol- lowing conclusions were noted: 1. As the tip Mach number was increased up to 0.86, the NACA 0012 rotor having -8° of twist encountered compressibility losses at progres- sively lower thrust coefficients. At tip Mach numbers of 0.91 and above, compressibility losses were evident at the lowest values of thrust obtained in the rotor tests. 12 NACA EM L58B19 2. No effects of compressibility were evident at any test thrust coefficient on the NACA 64-series rotor having -l6° of twist up to tip Mach numbers of 0.91- At tip Mach numbers of about 0.92 and above, compressibility losses were evident at the lowest values of thrust meas- ured in the rotor tests. 3- The measured power requirements of both rotors at the lowest tip Mach numbers are significantly more than predicted by usually reliable theory. This effect is believed to be due to the unusually large ratio of hub diameter to rotor diameter. The absolute values of the measured data are therefore not believed to be representative of full-scale rotors; however, the effects of Mach number on either of the rotors or compari- sons between rotors are believed to be valid. k. Calculations indicate that the drag-divergence Mach numbers of these rotors are about 0.1 higher than would be predicted from two- dimensional airfoil data. 5. For tip Mach numbers of 0.86 and above, the NACA 64-series rotor had higher figures of merit than the 0012 rotor for all values of thrust- coefficient — solidity ratio above 0.02. 6. As the tip Mach number increases, the maximum thrust-horsepower ratio is reduced. The optimum value for a given tip Mach number occurs at a higher disk loading as the tip Mach number is increased. 7. The maximum sound level measured in the tunnel test chamber at a point 21.5 feet and 140° from the vertical shaft on the NACA 64-series rotor was 122 decibels at a tip Mach number of 1.0. Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va., February 5> 1958- NACA RM L58B19 13 APPENDIX DERIVATION OF BLADE-TIP LOSS FACTORS The basic equation for tip loss (i.e., effective reduction in blade radius) due to three-dimensional flow at a propeller blade tip was devel- oped by Betz and is given in reference 9- In the terminology of the present report this equation becomes: 1.386 B = 1 — 1 + A 2 where A=^ aR If a uniform distribution of induced velocity is assumed, B=l-^86 If a triangular distribution of induced velocity is assumed, 2k NACA BM L58B19 A = 1.06 l/Crp or A« \/C T Since C T « 1.0 and I.386 ~ \J2, .4 B = 1 2C T NACA RM L58B19 15 REFERENCES 1. DeFrance, Smith J.: The N.A.C.A. Full-Scale Wind Tunnel. NACA Rep. U59, 1953- 2. Carpenter, Paul J.: Effects of Compressibility on the Performance of Two Full-Scale Helicopter Rotors. NACA Rep. IO78, 1952. (Supersedes NACA TN 2277.) 3. Powell, Robert D., Jr.: Compressibility Effects on a Hovering Heli- copter Rotor Having an NACA 0018 Root Airfoil Tapering to an NACA 0012 Tip Airfoil. NACA RM L57F26, 1957- k. Shivers, James P., and Carpenter, Paul J.: Experimental Investiga- tion on the Langley Helicopter Test Tower of Compressibility Effects on a Rotor Having NACA 632-OI5 Airfoil Sections. NACA TN 385O, 1956. 5. Gustafson, F. B.: The Application of Airfoil Studies to Helicopter Rotor Design. NACA TN 1812, I9I+9. 6. Gessow, Alfred, and Myers, Garry C, Jr.: Aerodynamics of the Heli- copter. The Macmillan Co., C.1952, pp. 72-73- 7. Rabbott, John P., Jr.: Static-Thrust Measurements of the Aerodynamic Loading on a Helicopter Rotor Blade. NACA TN 3688, 19 56. 8. Hicks, Chester W., and Hubbard, Harvey H. : Comparison of Sound Emission From Two-Blade, Four-Blade, and Seven-Blade Propellers. NACA TN I35I+, I9V7. 9. Glauert, H.: Airplane Propellers. Propellers of Highest Efficiency. Vol. IV of Aerodynamic Theory, div. L, ch. VII, sec. h, W. F. Durand, ed., Julius Springer (Berlin), 1935> PP- 261-266. lL NACA RM L58B19 TABLE I PHYSICAL CHARACTERISTICS OF NACA 0012 ANJJ NACA 61+-SERIES ROTOR BLADES Item NACA 0012 rotor NACA 64-series rotor Weight (blade alone), lb Radius, ft Chordwise center of gravity (blade alone), percent chord . . c e > in Tip chord, in Root chord (extrapolated to center - line rotation) , in Spanwise center of gravity (blade alone) , percent span Twist, deg a 7' (including hub) I (including hub), slug-ft^ . . . Airfoil section: Blade tip Blade root (extrapolated to renter-line rotation) .... 12.17 5.073 21+.19 8.98 6.16 16.00 1+2.60 -8 O.095 0A5 2.75 NACA 0012 NACA 0012 13-93 5.068 25.26 8.77 5-81+ 15.77 31.90 -16 O.092 0.1+0 2.88 NACA 61+ -006 NACA 6I+2-OI5 NACA RM L58B19 17 L-92293 Figure 1.- NACA 0012 airfoil rotor blade used in hovering tests. 18 NACA KM L58B19 ■a to n3 CD s c hO CO Q- J L •o o 4-> o CH 03 H O o <: o < 2 nJ J l_ cn an CO uo co i (1) H bfl t3 0) P o e p 01 w -P w 0) p u o p o 05 o bfl ■H 20 NACA RM L58B19 Figure h.- Sketch showing relative size of rotor head and blades. 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' X - \ X \ \ \ N ^ , \ \ % X n s X \ > \ N^ \ *> •^ XX V ^ > ^ ^ IW ■n sf 3 o CM o o •H o w a o 0) -P -P o cd a) O H rH -H a) O o en t3 to TJ 0) 0) -H 3 0) cd 1 S3 o o o -p • o ) — 1 1 0) I •H (in NACA RM L58B19 35 "r onq 17 <> C 008 007 ^ df 9^ / 005 c v- / 004 A / / 003 / ? — V— Experimental Calculated Condition Tip loss factor P A / / D Tip Mach relief b x i ~»-^t 001 O No tip Mach relief „ , 1 •«,« Jop/o p- Tip Mach relief B=l-i-38J ^T/2 — R- O — .0001 .0002 .0003 ' .0004 .0005 .0006 .0007 .0008 .0009 .0010 (b) M t = 0.95- Figure 15 • - Continued. 36 NACA RM L58B19 r 008 007 / 006 / if / 005 > S> / V 004 / 003 — 0- Experimental Calculated Condition Tip loss factor 0/ < £> t < / D Tip Mach relief B = 1 -_ <> No tip Mach relief . , QC V Tip Mach relief B»l-±*|se <2C T b • \/or/2 /H-Ct/2 r ■ < > .0001 .0002 .0003 .0004 .0005 .0006 .0007 .0008 .0009 .0010 CQ (c) M t = 1.0. Figure 15-- Concluded. NACA RM L58B19 37 1.0 -.6 / / / / / / / / / r / / c l, degrees 1 / / /' / / t^ / ', 1 ^y 3 — Ext] •apol ated r / \ \ a, d egres iS \ 4 X \ \ \ \ 2 \ -2 / / / \ V / / .3 .5 .6 Uach number .7 .9 1.0 Figure l6.- Two-dimensional NACA 6U-006 airfoil data used for theo- retical performance calculations. 38 NACA RM L58B19 a) .06 / / / /' / / / / / / / * .02 a , degrees 1 / / / / 4 A > / / ? — / 0,1/ — Extrapolated 1.0 .8 S a, d egre< !S / / \ -"*" - \ \ 5 / / .4 \ \ / / 3 \ \y .2 \ / 1 \ /' \ -1 / / \ -A .1 .2 .3 .5 .6 Mach number .7 .9 1.0 Figure 17.- Two-dimensional NACA 6U-OO8 airfoil data used for theo- retical performance calculations. NACA RM L58B19 39 10 00 CD • '1 • CO \ a CO • 4 • ■H- • 1/ / / V // V f • II ■P ' V \ \ \ ' \ ■ \ s C\ \ 1-1 o i-l 03 o CO o sr\o O CM O 00 CO OJ 01 fi +-' m • -P U \ •H I- A u <;> • B tH II B q; o I o -p o (U w 00 H 111 'aiJQJfl JO 9JrixTj 1+0 NACA EM L58B19 1 1 / LO / / K • / LO • O • r-K || sT \ \ \ \ N \ hs> \ Vs ^ O 00 o CO o o .3 u o -p o u ia CO S-. CD •5 o a CD CO J-, o 4 6 8 10 12 14 16 18 20 Disk Loading, lb / 2 Figure 21.- Effect of disk loading and Mach number on power loading of a 10-foot-diameter rotor having NACA 64- series airfoil sections and -16° of twist, a = 0.092. IIACA RI! L58B19 h3 sT U o •p o u Pa •H CO _p 0) h tQ M -P o 5 X! -P o w O C o •H "■8 •H I qp *8jnss9Jd punos NACA - Langley Field, Va. UNIVERSITY OF FLORIDA 3 1262 08106 585 5 -ERSITY OF FLORIDA DOCUMENTS DEPARTMENT 120 MARSTON SCIENCE LIBRARY P.O. BOX 1 1 701 1 GAINESVILLE. FL