7/27/2019 The Aircraft Engineer July 24, 1931

1/8

J u v 24, 1931 Supplem ent to FL I GHT

Edited by C. M. POULSENJuly 24, 1931

CONTENTSAir-Cooled Engine Power and Weight . By W. R. Andrews,A.F.R.Ae.S. . . - "Correction of Aeroplane Performance to Standard Atmosphere(Density Basis). ByClifford W.Tinson, F.R.Ae.S. , M.I .Ae.E. ...A Formula for the Buoyancy of the Wing Floats of Flying Boats andSingle Float Seaplanes. ' By A. R. CollinsTechnical Literature

49515556

which is of great concern to aircraf t and enginedesigners .W i t h a view to inves t igat ing a purely hypothet icalcase of an aircraf t us ing an engine of fixed b.h.p. andhaving a given range, the writer obta ined empir icalformulae for engine power and weight .AIR-COOLED ENGINE P O W E R AND W E I G H T .

B Y W. R. A N D R E W S , A . F . R . A e . S .Mr. Andrews, who is on the Technical Staff of A. V.Boe dc Co., Ltd., has previously contributed articles to

THE AIRCRAFT ENGIN EER. In the present article heturns his attention to the question of thrust horse-power available at the airscrew for every pound ofengine and fuel. Emp irical formula} have been evolved,which appear to give at least sufficient accuracy forpreliminary investigations, Mr. Andrews basing hisforrnulce on the assumption that power output is pro-portional to cylinder capacity.THE question of the thrust horse-power available atthe airscrew for every pound of engine and fuel is one

40 5350

ui2(3 -Z "

1 ,';ZD WC "6.

TOTAL EN81**" VX, J , 300 iSO

HE CA \CsT Y- (CUBIC !NS.)

These formulae are sufficiently accurate over a largerange of powers to be of use to others in teres ted insimilar problems.As pointed out, these relationships are only empirical,and their only justif ication is that they agree with theactual conditions within limits which are not too l a rge .The varying factors in engine design, such as c rank -shaf t r .p .m. , number of cylinders , capacity of cylinders ,mater ia ls used , etc., would at first sight make the t a s kof generalis ing almost hopeless .However , it is obvious that s ince cylinder capacity issomewhat limited by various considerations, as is alsothe compression ratio, then the power output is pro-por t ional to total cylinder capacity at the satne design,r .p .m.726a

7/27/2019 The Aircraft Engineer July 24, 1931

2/8

SUPPLEMENT TO[FLIGHT

Ma ke r .Gnome and Rhone

Lorra ine

Saimson

SiemensWalter

Armgtrong-Siddeley

ElizaldeFiatKenaultHiapano-SulzaS.F.F.A

Engine.Name.

Titan II; 5 BC .. oKi-Kdr . . Kb . .10 0 h .p110-h.p120-h.pMizarAlgolAntarps3 A D6 AC7 AC9 AD9 AC9 N - C T9 A B18 ADS.H. 13 a8.H. 2q60 NZVega IKoculus85 NZVenus I13 0 NZMars IGenetGenet MMongooseGenet MLynxLynx MJ a g u a rLeopardDouble M ongooseDragon VDragon V IISuper-DragonA. 50

9Q" B "" A "

TH E

No . ofCylinders.

5575557914357999918595557

I555777141410579779937

50A IR C R A FT

TABLE I

Bore.In .

5-755-755-754-924-925 1 25-515-615-512-753-943-942-763-963-964-924-924-136-064-134-135-324-134 1 35-324 1 34 1 3404-255- 04-25505-255 06-05-05-12fi-95- 93-943-944-925 -04-144 1 4

Stroke.In .

6-56-56-55-515-515-515-95-95-93-395-125 1 23-385-125-516-77 0 84-727 -44-724-726 -34-724-726-74-724-724 -04 -55- 54 -55- 55- 55- 57 -55- 55-515- 97-484-734-735- 95-514-934-93

ENGINEER

Capaci ty,Cub. in .

84 284 21,18052 152 156 698 51,2641,9706031 243 618256 761 01,1452,421

31 61,92031 631 672 044 244 21,04056 956 925 131 853944 775 583 21,5122,9701,08056 51,1291,8404 0 340 31,0099 7319 946 5

NormalR .P .M.= N .1,8002,0001,9501.3501,6501,7001,8001,8001,9002,0001 80 01,8002,0001,8001,8001,7001,7001,8501,6501,4001,7501,8001,4001,7501 75 01,6001,7502,2002,2001,8502,2001,9002,0001,7001,8002,2001,8001,8001,8001,8002,0001,8202,000

N orma lB . H . P .

25 528 537 010011 012 024 030 050 0

1660954612 015 023 050 07542 0607518 58511 024 013 014 58210 315514 022 026 041 674 0342

16 532 062 51 0 010 02502S0

PC.N'7'

0-001480-001540-001450-001150-001100-001080-001190-001160 0 0 1 1 90-001200 000940-001060-001150001030-001200-0O1030-001060-001140-001140-001110-001350-001250-001130-001240-001150-001220-001280-001380-001370-001440-001320-001370-001410-001400-001300-001340-001420-001380-001390-001210-001130-00120-00117

JU LY 24, 1931

Weight .Lb .= W .

56252 862 734 434 835458 365 293 0

7524 228 515 437 430 858 399 024 289 122 522 742 028 029 155036 535 2

37 566 084 02 8 630 858 359 816 529 7

P-76:;

47243642050053151451548 648 367462050851056038 351448 151749049 944045047 846047 948 145 1

44 246 941 14 9 156 549 354 5

The da ta used in the inves t iga t ion i s suppl ied byCa pt. Swan 121 his review of th e Pa r is showsee An t-CBAJT ENGINEER, February 27 , 1931.At f i r s t s ight the re la t ionship be tween capac i ty andr.p.m. i s not apparent , but by dividing b .h .p . by capa-c i ty and plot t ing the re sul t aga ins t r .p .m. i t was foundtha t for unsupercharged a i r -cooled engines

P = K C (1)where P = norm al b .h .p .C = capac i ty of engine , cub. ins .N = r .p .m. a t normal speed.K = c ons t a n t .The effect of varying compression rat io is neglected,especially as an increase in compression rat io f rom 5 to5.5 only means about 3 per cent , increase in power.A var iat ion of 3 per cent , is obviously outside theaccuracy of the formula.The con stan t K var ies tor the different de sign condi-

t ions adop ted by each f irm, but for any pa rt i cu lar f irmthe va r ia t ion for the i r r ange i s smal l .Table I g ives the va lue of K and the pr inc ipa l da taext rac ted f rom Ca pt . Swa n's a r t ic le . The enginesse lec ted a re ungeared and unsupercharged.The biggest range of engine power is provided by theArmstrong-Siddeley engines, varying as i t does f rom82 b.h.p. for the Genet to 740 b.h.p. for the Leopard.Even over this large range the value of K only var iesbetween a minimum 0.00132 and a maximum of 6.00144.The mean value of K for the Armstrong-Siddeleyengines i s 0 .00138, so tha t the maximum devia t ion ofes t imated power f rom the ac tua l i s something le ss than 5 per cen t .

In a few isolated cases the constant for some par-t icula r engine va r ie s r a the r a lot f rom the mean va luefor the range to which it belongs.The most notable example is supplied by the Saimsonengine 5 AC, which has a value for K of 0.00094,whereas the mean for the Saimson range is 0.00107.This r epresents a devia t ion of nea r ly 20 pe r cent . ,a l though the remainder of the range i s f a i r ly cons is tent .The ac tua l meaning of the constant K is a l i t t l eobscure , and is probably greatly inf luenced by the speedof the induc t ion gases a t the des ign speed.One thing i s ce r ta in , tha t the constant i s no indica -t ion of the thermal eff ic iency of the engine; nor is i tany indica t ion a t a l l a s to the qua l i ty of an enginer unn ing in a t h r o t t l e d c ond i t ion .We can now pass on to the problem of engine we ight .I t i s r a the r surpr is ing tha t i t i s a t a l l poss ible togenera l i se on this subjec t .F ig . 1 , however , shows tha t engine we ights a re mainlydependent upon the tota l cyl inder capac i ty .The law for the middle curve drawn isW = 2.93 C-7li:i . .where W = weight in lbs . (2)

The upper and lower curves show a va r ia t ion of 10 pe r cent , f rom the mean curve .The notable except ion to the curve i s the Sa lmson9 -N-CT eng ine, which, if the p ublished d at a is corre ct ,weighs 66 lb. less than 9 AC engine, which has the samebore but shor te r s t roke than the 9 N-CT engine .No genera l i sed re la t ionships can account for suchcont ingenc ies .726b

7/27/2019 The Aircraft Engineer July 24, 1931

3/8

24 , 1931 61THE AIRCRAFT ENGINEER SUPPLEMENT TOFLIGHTBy combining equations 1 and 2 it is possible to findtfc- law of engine weight in terms of b.h.p. and r.p.m.Rewriting 1 gives

C = KN"1Substituting this in 2 gives

W = 2.93 ["2-93

(3)

Since it has been shown that, although K varies withdifferent makes of engines, it is practically constant forany partic ular mak e. W ithou t loss of accuracy, the re-fore, we may write;where K2 should also be constant for any particularmake.Reference to Table I will show that within fairly closelimits this is the case for all normal engines.The weight per b.h.p. of any engine can then beexpressed as

weight/b.h.p. = - ^ J T ^ ; (5)Although this article is confined to ungeared engines,mainly on the ground of insufficient data, it is concludedthat the addition of gearing adds from 10 to 15 percent, to the engine weight.By designing an engine to run at 1,400 r.p.m. instead

of 2,100 r.p.m., but giving the same b.h.p., would add25 per cent, to the weight.The saving in weight due to fitting gearing of 1 to.667 ratio is from 10 to 15 per cent, of the weight ofthe engine designed to run at the airscrew speed.

CORRECTION OF AERO PLANE PERFORM ANCE TOSTANDARD ATMOSPHERE (DENSITY BASIS) .By CLIFFORD W. TINSON, F.R.Ae.S., M.I.Ae.E.

In subm itting the following article for publication inTHE AIRCRAFT ENGINEER, Mr. Tinson, leho is a memberof the technical staff of the Bristol Aeroplane Co.,Ltd., states that he fears the article is not very " high-brow," but that, his " excuse " for writing it is that hehas found that the standard atmosphere is apt to berather confusing, and he wished to be able to delegatethe performance corrections to junior members of thestaff.The following article describes a m ethod of correctingexperimental test-flight report figures to standardatmosphere, in which the aim has been to obtainaccurate results in the shortest possible time. Thestandard atmosphere has been taken from the Ameri-ca n N.A.C.A. Report, No. 218, " Standard AtmosphereTables and D ata," and the formula for obtaining thetime to height from the rate of climb from E. & M.

No. 1,316.Mr. Tinson does not claim that the article containsanything new , or anything that differs materially fromthe methods commo nly employed in performance cor-rections. He has, however, made an attempt to setdown the principles involved in a general way, and adescription of the work as it is carried out, keepingthe article in sirnple language for the benefit of anyonewho does not yet understand the procedure very clearly.INTRODUCTION.

THE density of the atmosphere, as of any gas, in-creases with an increase of pressure, and decreases withan increase of temperature.

It may be generally stated that the pressure and thetemperature both decrease as the altitude increases,but the nett effect is a reduction in density as the alti-tude increases.Owing to climatic conditions, the pressure and tem-perature vary from day to day, and consequently thedensity is not always the same.The performance of an aeroplane depends on thedensity of the air because the power output of a nor-mally aspirated (or non-supercharged) engine decreasesas the altitude increases. This is due to the fact th ata smaller mass is aspirated on the induction stroke asthe atmosphere becomes less dense.A certain minimum horse-power is required in orderto maintain level flight, the speed at which flight ismaintained under these conditions being fixed by theaerodynamic chara cteristics of the aeroplan e. Anychange from these conditions, such, as the change ofincidence of the wings, demands more horse-power, orthe aeroplane would lose height.If the horse-power is increased by opening the throttleof the engine, the horse-power available in excess ofthe aerodynamic minimum permits an alteration to wingincidence without loss of height, and thereby endowsthe aeroplane with a range of speeds over which levelflight is possible.When the attitude of the aeroplane is such that thewings are at the incidence corresponding to minimumhorse-power required, but with the throttle opened, theavailable horse-power in excess of that required to keepthe aeroplane in level flight is translated into verticalforce, thus giving the aeroplane the ability to climb.

When the aeroplane has climbed to an altitude atwhich the horse-power has decreased to a value suchthat the entire horse-power is devoted to maintainingit in level night, then the aeroplane is said to havereached its absolute ceiling.The absolute ceiling may be mathematically definedwith precision, for it is the altitude at which the rateof climb is zero with full thr ot tle . Owing to the extrem elength of time taken to climb the final two or threehundred feet, it is usual to adopt a practical ceilingknown as the " Service Ceiling," and this is defined asthe altitude at which the rate of climb at full throttlehas fallen to 100 ft. per minute.Variations in atmospheric density not only affect therate of climb through absolute units of altitude abovethe aerodrome, but also affect the height of the ceiling,and it is therefore impossible to compare the perform-ances of different flights unless the figures have beenreduced to a comparative basis.For example, if an aeroplane is imagined to haveso little power that its absolute ceiling is only 1,000 ft.(indicated) above the level of the aerodrome on a daywhen the temperature and pressure gave a density valueof p, and that the next day the density at this aero-drome changes to />,, on the second day this aeroplanewould be capable of reaching, say, 1,100 ft., an appa-rent increase of 10 per cent, in ceiling.Its speed range at the aerodrome level would also beslightly increased on the second day.

The percentage difference in performance is generallymuch less than that given above, which is only intendedto illus trat e the effect of density changes. Aerodromes,however, are to be found at all kinds of different heightsfrom sea level, so it will be understood that the per-formance of aeroplanes can only be fairly compared byreducing the observed figures to a definite standard.The process of reduction is known as the Correctionof Performance.After making careful investigation into atmosphericconditions, varying altitudes, and the different seasonsof the year, the National Advisory Committee for Aero-nautics of America recommended the adoption of thefollowing values as representative of the StandardAtmosphere, and these were adopted in January, 1925.726c

7/27/2019 The Aircraft Engineer July 24, 1931

4/8

52SUPPLEMENT TOF L I G H T THE AIRCRAFT ENGINEER JULY 24, 193!S T A N D A R D A T M O S P H E R E .Standard pressure p0 to be the pressure exerted by acolumn of mercury 760 mm. (29.921 in.) in height.Standard temperature t0 = + 1 5 deg. C. = 59 deg. F .Standard absolute temperature To = 288 deg. C. =51S.4 deg. F.Standard specific weight gP = 1.2255 kg. /m.3 = 0.07651lb./ft . sStandard gravity g = 9.80665 m./s ec.2 = 32.174 ft./sec*Standard density P = 0.12497 = 0.002378.Standard temperature gradient a = 0.0065 deg. C./m. =0.003566 deg. F./ft.The Standard conversion factors ar e: 1 metre = 39.37 in. = 3,280833 ft.1 kilogram = 2.204622 lb.

Specific weight of mercury at 0 deg. C. - 13595.1kg. /m. s = 848.7149 lb./ft.3The above values are taken from the NationalAdvisory Committee for Aeronautics Report No. 218.It is necessary, then, to correct the observed perform-ance figures of an aeroplane to reduce them to whatthey would have been had the condition of the atmos-phere been " standard " all the way up to ceiling.To enable this to be done, it is obviously necessaryto carry a thermometer on the aeroplane in some posi-tion where it will record the general temperature ofthe surrounding air (not in the cockpit), and to carryan altimeter with a dial locked to show zero feet ormetres altitude when the temperature and pressure arestan dard . Initia lly, this altimeter must have been putinto an exhausting box at a tem peratu re of + 15 deg.C , in which the pressure of the air is reduced to760 mm. by pump, and then locked.It is important to note that performance cannot bereduced to standard atmosphere unless an altimetersuch as this, calibrated in an exhausting box and locked,is carried. An altimeter is only an aneroid barom eter,or pressure indicator, but is marked in thousands offeet for the convenience of the pilot.For scientific purposes the indications of the ordinaryaircraft altimeter with movable dial mean nothing, asis obvious, for, when the pilot sets the dial to read zeroheight on starting, the aerodrome itself may be severalhundreds of feet above sea level. By calibrating thealtimeter as described and locking it, it is transformedinto a pure pressure indicator, the zero of which meansthat the pressure is equal to 760 mm. of mercury.The air speed indicator, like the altimeter, measuresthe pressure of the air, but is graduated in miles orkilometres per hour for convenience.There is a definite relation between the pressure andthe speed of flight, and the instrument is calibrated sothat it will indicate the correct velocitj- at ground level,tha t is, at " standa rd " ground level. At any othervalues of temperature and pressure the velocity indi-cated by thfr ins trum ent is not tru e, and must becorrected.In other words, when the aeroplane is high up andthe air is of much lower density, the pressure in theair speed indicator piping is insufficient to displace thediaphragm which actuates the needle to the extentnecessary to indicate the true velocity.The pressure in the pitot tube is proportional to thedensity of the air and to the square of the speed of thepitot tube through the air, so that it is necessary todivide the indicated air speed by A / - where is the

V Po Prelative density, in order to obtain the correct speed.The relative density is obtained by dividing densityat height by density at zero standard feet; these figuresmay be taken from tables in the N.A.C.A. ReportNo. 218, quoted above. The relative den sity at 20,000standard feet height is 0-5327, so that an indicated air

speed of 130 miles per hour at this height would correctto 178 miles per hour.At 30,000 standard feet height the density falls to0-3740, so here, if the true speed were 150 miles perhour, the air speed in dic ato r would show only 91-8 niilesper hour.In addition to this correction for air speed indicatorreading, a further correction should be applied whichmay be im po rta nt. The air speed indicat or will almostcertainly be subject to a " position error," due to theproximity of the pitot head to the structure of theaeroplane. As a rule, the instrum ent in the cockpitgives a lower speed than it should, because of the reduc-tion of velocity at the pitot head due to the strut towhich the latter is attached, and to the effect of thewings on the air near the pitot head.There is no method at present whereby the positionerror may be calculated, and to find the magnitude ofthe correction experimentally, it is necessary to carrya second pitot and indicator, the pitot being eithersuspended below the aeroplane or mounted on a longhorizo ntal m ast pr ojecting a considerable distanceahead of the wings, so that it travels in air undisturbedby the aeroplane.Usually, the indicated air speed as stated in the pilot'srepo rt is stated ag ainst altime ter height. It is neces-sary, therefore, to correct the height to standard atmos-phere at the same time as correcting the speed fordensity, and to draw a speed curve so that the speedsmay be finally quoted against regular intervals ofstandard height, such as 15,000, 20,000, 25,000 ft., etc.When a machine is being put through performancetrials, it is usual to carry a recording barograph and arecording air speed indicator. These instrum ents pro-vide graphic evidence of the performance in additionto the records taken by the pilot.From the nature of the records made by these instru-ments, the data available are not generally of a higherdegree of accuracy than that provided by the pilot'sobservations, assuming, of course, that the pilot'sfigures for altitu de have been tak en from a calibratedand locked altimeter and not from an ordinary alti-meter, and provided that the air speed indicatorif nota specially calibrated instrum ent for test flyingh asbeen calibrated in the usual manner by means of acolumn of liquid in a U-tube.A greater degree of accuracy may be obtainable fromthe recording instruments when facilities are availablefor calibrating them with proper laboratory apparatus.As a rule, the height or speed and time scales whichare printed on the sheets of paper used to wrap roundthe revolving cylinders of these instruments are fairlyaccurate in themselves, but are difficult to sub-divide,and the mark made by the recording pen is too broadfor a delicate analysis to be made. Fu rth er , unless theapparatus is in the charge of a responsible person, thereis no certainty that the base line of the printed sheetis truly horizontal with respect to the spindle of thecylinder.For these reasons it is generally advisable to considerthe recording instruments as independent witnesses onlythat the required height has in fact been achieved, andthat the required speeds have been taken whpn themachine has been flyiig level (i.e., not climbing ordiving) for a reasonable length of time, and are notinstantaneous readings of the air speed indicator.W ith this evidence the pilo t's recorded observationsmay be used with confidence for correction of perform-ance to standard atmosphere conditions.

CORRECTION O F C L I M B .The data required are as follow:

(a) Height from locked altimeter, feet;(b) Time to altimeter height:(c) Air temperature at altimeter height.

7/27/2019 The Aircraft Engineer July 24, 1931

5/8

24 , 1931 53THE AIRCRAFT ENGINEER SUPPLEMENT TOF L I G H TEXPERIMENTAL. "TEST FLISHT TCPOgT Ng 694

WCK3MT AS T EST ED :

GIKUNO TE M PCPATuW:

34^4-.X&S3 INSi^ *C

QBMARKfl-.. _BAl_L.AT CA(TITID if^_i_lUOW M O V A B U C _T*!t t_T*5 rX_ lro*o_ . ^ c ? ^ . T _ i T ? ? l * .ITACK ^ GBNC

FIG. IO-lMft

a I.C.A M

ALTlMBTEItHCraHT, FT

+ soocoo4OOOfeOOOBOOO

IZOOO14000i&OOOITOOOIOOOOI9OOO20OOO2IOOO8EOOOB3OOO24OOO3COOO

CTIME

M I N S -eece-

o esZ 13S Z3A 3?

T OSe 37

io 3oit SoIS. 3K13 -*3lA Sftia soIS IS

2o OSSB COSf i OS

MINUTEO

eessS'ZIT

A>i>

7-O63a bi&

IO6OO1 1 SOOIE-S33

i4-essKb-SOO18BOOCOOQ9ZE -333Eaooa

d eTEMP1WATURC C

oaeetTVEo

i CT

3,

- 3- 7-

-- i s-i & 1 9

-- e s

- e

FAIRED! O( S O

7 - O3 d2 - 7i O

- s o-7-O

-to x- i i -e- I 3 Z-1-4 7

IB-7zoo-32 IOSI4O

49OO

s

ORto76Z6

7I7SfeSfeO

6PC1OCOPffECTtDFOB" O*tsmiT>M.*M

tT6BIT46tfct, a1 *1 274z-szeS-I9O

VTWUIAIR

M F H1 Tfc E&4ITfc O^Ci!&6'6ESic>e -i 9o

urSPEED ilM

STANDARD AT**OG**nett&17fc- 21 M. ^. H A.7" IOOOO FEBIT174-6 I6OOO171-5 EOOOO

I t isalso usetul to know theengine revolutions duringthe climb, asthese, plot ted onthe final cha rt , providemost useful information for record andcomparison.All theabove isgiven inthepilot 's test fl ight reportand should beentered up in a table such asthe onei l lustrated (Fig. 1).In this table it will benoted tha t there are twocolumns fort e m p e r a t u r e , oneforobserved figures andthe other termed " fa i red . "The pilot maybeunde r thedisadvantage tha t hecannot read the thermometer with such accuracy as hecan read theother ins t ruments , norpe rhaps at theexact moment when thealt imeter needle crosses an alt i-tude level, and asit isrequi red toknow the t e m p e r a t u r efairly accurately to find theequivalen t s tandard height ,any slight discrepancies maybesmoothed outbyplot-t ing theobserved figures and then drawing a fair curvethrough thespots, as shown inFig. 2. This line shouldnot vary more than about half a degree ei ther way,however, as it isreasonable toassume tha t the pi lo t ' sfigures are within that tolerance.F i g . 2shows the fa i r ing ofthe pilot 's observed tem-pera tures , and ind ica tes tha t the fa i r ing is res t r ic tedto the l imits ofpossible observational error Thecurveis not faired from endtoend, as it were .In passing, anysharp i r regular i ty inthe slope of thet empera ture curve , such as is i l lus t ra ted , is general lyaccompanied bywaves inthet ime toclimb curve . Theseusually occur when there arevert ica l curren ts inthe airwhich are set up bytemperature differences due toclouds,etc.

There maybe a t emp ta t i on tosmooth out the t imeto climb curve when this is the case, butbearing inmind tha t the purpose ofthe work is tot ranspose thet imes inal t imeter he ight tot imes ins tandard heightonly, andnot to tryand make a badly-shaped climbcurve into agood-looking one, it isadvisable tostrict lyconfine anyadjustments made byfairing curves to thel imits ofobservational error which apilot could reason-ably claim.These " faired " t empera tu re s may then be drawnona curve constructed from N.A.C.A. Report No. 218

s\ m -C----O--{*

-

FIG.2 1

y

12%e

7/27/2019 The Aircraft Engineer July 24, 1931

6/8

SUPPL E ME NTTOF L I G H T54

THE AIRCRAFT ENGINEER J U L Y 24, 1931

connect ing the a l t imeter reading , the t emperature andthe equivalen t s tand ard height . Then ateach a l t imeterheight the equivalen t s tandard height may beread offag a i n s t theintersect ion of t em p era t u re w i t h a l t i m e t e rheight .The s tandard heights corresponding toal t imeter read-ings may then be entered inthe t ab le . It is preferable ,however, toobtain the s tandard height wi th ag r e a t e rdegree ofaccuracy, and the fol lowing equation connect-ing a l t imeter height , t emperature and s tandard heightm ay beused for alllevels below the lower level of th ei so thermal a tmosphere , viz., below 35,332 ft., or10,769 m .:

H , =(1.238 H a +120 to -1,800).H s =S t an d ard h e i g h t infeet .H a =Alt imeter height infeet (locked al t imeter).to Observed temperature indegrees C.(The plus s ign becomes minus, of course, ift0is nega-t ive . )Having fi l led in theequivalen t s tandard heightsinthe t ab le , theclimb ins tandard a tmosphere isknownfrom thet ime column, but thecommencement ofth i scl imb isvery rarely from zero stan dar d feet . This couldonly be soifthe barometer reading at the aerodromelevel was 760 mm. and the g ro u n d t em p era t u res+ 15 deg. C.For ins tance, the commencement of the c l imb mighthave taken place when thelocked al t imeter indicated20 0 ft. Ifthe t emperature had been 4- 16 deg. thestart would have been made from aheight of+427.6 ft.in s tandard a tmosphere . Ifthe t emperature had been1 2 | deg. instead, the start would have been made from aheight of-52.4 ft. ins tandard a tmosphere .

RATE OF CLIMB .The average ra te ofclimb atvar ious a l t i tudes isveryeasily calculated from the t ime cl imbs tostandard

heights , but onplo t t ing theresu l t s itwill usually befound that thepoints arevery scat tered , nomatterhow careful ly the previous work has been carried out .I t isdifficult then tod e t e rm i n e at s ight aline whichrepresents the ra te ofcl imband, therefore, the cei l ingaccurate ly . Afewt r i a l r a t e ofcl imb curves willprobably benecessary before oneisfound which, onworking back, gives thet ime toclimb figures whichagree wi th the t ime curve.In the case ofanormal ly asp i ra ted engine, the ra teof cl imb curve isusual ly as t r a i g h t l i n e , t h a t is to Bay,the reduct ion ofrate from ground level toabsolutecei l ing isregu lar. Such an engine , however, may possessfeatures which permi t par t i a l main tenance ofgroundpower forthe fi rs t few thousands of feet , and insuchcases itisdesirable tocheck back, after having deter-mined ar a t e ofcl imb curve which gives the correct t imeto aheight near the cei l ing , tover i fy that it also givesthe correct t imes at heights nearer the ground.In the case ofasupercharged engine, the ra te ofclimbis usual ly main ta ined , oreven increased, from theground up toacer ta in a l t i tude, af ter which itfalls inthe normal manner . Insome cases, the rate ofclimb isnot main ta ined , bu t the fa l l inr a t e is not somarkedfor the fi rs t few thousands of feet .The shape of the curve below the al t i tude where thefall-off commences to benormal depends on the type ofsupercharger , e tc . , and isdifficult todetermine accur-ately when the plot ted points are widely scat tered. Thetime toth e altitu de wh ere th e fall-off commences to be

T H E nwe -4EceAgyTOCUIHB FWOM renoSTAMDMPHE MgKMTATWHCH.ri-lOMT tCTUAoLY COMMENCED

From th i s example itisclear that the curve obta inedby plot t ing cl imb t imes against the figures fors t an d a rdheight mus t bemoved across thep ap er as showninF i g . 3, so th at th e curve intersec ts zero heigh t , which isthe same as adding (or subt ract ing , asthe case may be)the t ime necessary to cl imb through the number ofs t an -dard feet separat ing zero standard height from thepoint where the cl imb actual ly commenced.I t ismore convenient , of course, tos t a t e thecl imbtimes finally ast imes toregular in terval s of s t an d a rdheight . These may then be read off thecurve andentered inthe appropr ia te co lumn inthe t ab le .A t thefoot ofthis column arespaces inwhichtoenter service andabsolute cei l ings.These values are obtainable byplo t t ing the "r a t e ofcl imb "cu rv e .

normal can be found fairly accurately, andas thein terval t imes to al t i tudes below this level areno t 01very great in teres t , the preci se shape ofthe curve belowthis" al t i tude, that is,wh et h e r itisas t ra ight l ine ornot , isre la t ively unimportan t .For th i s reason , it isadvisable tocommence by assum-i n g t h a t ther a t e ofcl imb curve iscomposed oftw os t ra ight l ines , the one commencing at the ground andgoing uptowhat may be called the"superchargel i m i t , " an d theother commencing atthesuperchargel imi t and going up tocei l ing.The height of the supercharge l imi t can befixedtorthe prel iminary curve by inspect ion ofthe engine powercurve ifavai lable; fai l ing this , the r .p.m. cl imbing wil lgive afa i r ind icat ion .A l ine mus t then be drawn throug h the p lo t ted poin ts

726 /

7/27/2019 The Aircraft Engineer July 24, 1931

7/8

TULY 24, 1931 THE AIRCRAFT ENGINEER SUPPLEMENT TOFLIGHT

i j.SOOfc

sooo

\

Fl(3 4

\

of rate of climb, terminating at the supercharge limit,and from it the time to supercharge limit calculated andcompared with the ac tual time take n. If agreem ent isnot achieved, the line must be adjusted until its inclina-tion and upward extremity are correct.Above the supercharge limit, the reduction of rate ofclimb may be first considered regular, and a trial linerun through the plotted points is drawn and the timestaken off it as before for comparison with actual times.(To be concluded.)

A FORMULA FOR T HE BUOYANCY OF THE WINGFLOATS OF FLYING BOAT S AND SINGLE FLOATSEAPLANES.

BY A. R. COLLINS.The hull and aerostructure of a flying boat, or themain float and aerostructure of a single-float seaplanehas, in general, a negative transverse metacentricheight in the upright condition so that, in the trans-verse direction, it is inhere ntly unsta ble. In thiscountry, the conventional method of providing positivetransverse stability is by fitting wing floats, and thefollowing note deals wiih a simple formula for obtainingthe desirable buoyancy of such floats.Let us first consider the function of a wing float,

and how it gives to the aircraft a virtual positive meta-centric height in the transv erse direction . Refe rring tothe sketch, letW = the all-up w eight of the fly ing boat or single-float seaplane.w = the total buoyancy of one wing float.aft = the water-line in the upright condition.CD = the water-line when one wing float is com-pletely submerged.

0 = the angla of heel or roll to submerge completelyone wing float.G = the centre of gra vity of the complete aircr aft.B and B 1 = the centres of buoyancy of the hull or mainfloat (with its aerostructure) in the uprightand inclined conditions, respectively.M = the transverse metacentre in the uprig ht con-dition, i.e., the point where a vertical linethrough B 1 intersects the line joining B to G.6M = the transverse metacentric height of the com-plete aircraft in the upright condition.d = the distance from the centre line of the hull ormain float to the centre line of a wing float.

Since M is below G, the aircraft has a negative GM inthe upright condition, which gives rise to an upsettingnioment when the hull is displaced from the vertical.As BOOH as one wing float touches the water, a rightingttoment is called into play, which gradually overcomes

the upsetting moment of the hull and aerostructure,and gives to it a virtual positive metacentric height.Now when one win^ float is completely submerged,the upsetting moment of the hull or main float (with itsaerostructure) = W.GZ (where GZ is the perpen-dicular from G on to B'Mproduced)= W.GM. Sin 0.= H (say)The righting moment of the wing float= w.d. Co s 0= F (say).'. The reserve righting moment of the wing float= F - H.Hence the ratio of the reserve righting moment of thewing float to the upsetting moment of the hull or mainfloat

= - ^ = R (say) F H = R.H= R (W.GM. Sin 0)= W (11.GM) Sin 0.i.e., when one wing float is submerged completely, itgives to the aircraft a virtual positive inetaccntricheight = B.GM.Now a good empirical rule for the transverse meta-centric height of a twin float seaplane is" Transverse m etacen tric heigh t in feet = "v'Wwhere W is the all-up weight of the seaplane inpounds."

Hence , to make the stability of a flying boat orsingle-float seaplane consistent with good twin-float sea-plane practice, the virtual positive metacentric heightin feet, with one wing float just submerged should= *\/W . In other words, one should aim at the samemeasure of positive transverse static stability for flyingboats and single-float seaplanes, as has been foundadequate and desirable for twin-float seaplanes.Now let h = the negative metacentric height of thecomplete aircraft in the upright condition. \h will bepractically constant for values of 0 up to about 10,which will cover the range of angles of heel or roll, tosubmerge completely one wing float, for most flyingboats and single-float seaplanes.]

When one wing float is just submerged we have(using the same notation as before) upsetting momentof hull or main float (with its aerostructure)= W.fc. Sin 6.Righting moment of the wing float = w.d. Cos 6..'. Reserve righting moment of the wing float= w.d. Cos 6 - W./i. Sin 0..'. The ratio of the reserve righting moment of the wingfloat to the upsetting moment of th e hull or main floatw.d. Cos 8 - W.h. Sin 6= W.ft. Tw.d. Cot 6

W.fc. - 1 = R (say).Now it has already been shown that the virtual positivemetacentr ic height of the complete a i rcraf t with a wingfloat ju st subm erged = R.Ti. an d this should be = '/W

7/27/2019 The Aircraft Engineer July 24, 1931

8/8

SUPPLEMENT TOFLIGHT JU LY 24, 1931TH E AIRCRAFT ENGINEERj'w.d. Cot 6[ WJw.d. Cot 6

W

-l\h=

.-. w.d. Cot 6 = Wh + '= W(h + w = W: Tan 6 (h +

Large p ositive pitching mom ents due to the floats were found a t high a; ginof incidence . W ith eleva tors hard u p ( 34-1) the seaplane trinmuT; atabo ut 49 incidence as compared with 23 -5 incidence when the floate vertremoved (Fig. 14). Of the to tal pitching mom ent due to floats, about one-tnta]is accounted for by interference of floats on tailplan e. The floats also reducethe elevator control in the region of incidence 3555" (Fig. 15).The spinning calculations, based on the results of the rolling and Bideslinexperiments on the model with and w ithout floats, do not indicate that ther*should be any difficulty in recovery from spinning for the Fairey HIF s e a .plane, although the margin of safety in spinning is apparently less than forthe corresponding land machine. The nearest approach to danger wouldappear to be for a steep spin at an incidence not greatly in excess of thestalling angle, but it is in this region that the results of the model experi-ments may be most liable to scale effect.

which gives us a formula for the total buoyancy of onewing float wherew = total buoyancy of one wing float in pounds.W = the all-up weight of the seaplane in pounds.6 = th e angle of roll or heel, in degrees, to submergecompletely one wing float.h = the negative metacentric height (transverse) ofth e complete aircraft, in the upright condi-tion, in feet. [It should be noted that " h "will be numerically positive in the aboveformula.]d = distance , in feet, from th e centre line of thehull or main float to the centre line of thewing float.I t should be noted that considerations of height ofwing structure and span and area of wings are factorsthat should be taken into account in assessing th erequirements of the wing float buoyancy for particularaircraft, but the above formula will be found to givesatisfactory results lor the average conventional types

of flying boats an d single-float seaplanes.In the case of small single-float seaplanes (say up to2,000 lb . all-up weight) operating in sheltered waters,th e buoyancy of a wing float given by the above formulamay, perhaps, be reduced by about 10 to 20 per cent.,depending upon th e all-up weight.In the case of a flying boat or single-float seaplanehaving a small positive metaoentric height in the up-right condition, th e above formula is still applicable butin this case " h " will be numerically negative whensubstituted in the formula.

TECHNICAL LITERATURESUMMARIES OF AERONAUTICAL RESEARCH

COMMITTEE REPORTSThese Reports are published by His Majesty's StationeryOffice, London, and may be purohased directly from H.M.Stationery Office at the following addresses : Adastral House,Kingsway, W.C.2; 120, George Street, E d i n b u r g h ; YorkStreet, Manchester; 1, St. Andrew's Crescent, Cardiff; 15 ,Donegall Square West, Belfast; or through an y Bookseller.

SPINN ING OF A MODEL OF THK FA IR EY I I I F . SEAPLANE.By H. B. I rv ing , B.Sc . , and A. S . Ba t son , B.Sc.R . & M. No . 1356 (A e. 487). (15 pages and 27diagrams.) June , 1930 . Price Is . ne t .The object of th e experiments wag to prov ide data bearing on the spinningproperties of a twin float seaplane and to Investigate the effect of variousmodifications of tailplane on the moments given by the tall in a spin ; alsoto find th e effect or differential an d floating aileron s and " interc epto rs "In a spin.Measurement of rolling and yawing moments due to rolling were madeon the model with and without fin and rudder (incidence 22-4 to 608).The fln and rudder gave a moment opposed to the motion at all but thehighest incidence, when the moment became positive and reached a valueof Ip x 103 - 4-6 .Of the three forms of aileron control tried, the floating ailerons appear tooffer the grea test Improve men t as regard s recovery from a flat spin . Differ-ential ailerons are most beneficial at the lowest Incidence (42-4) bu t giveless rolling moment (body axes) at 60-9" than ordinary ailerons, with, how-ever, some reduction of yawing mom ent (Figs. 1012). Similarly, theeffect o t ' Intercep tors " falls off as th e In cidence Increases and Is small

DETONATION, MIN ER A L LTJRIUCATING OIL S A ND BLENDEDFUELS. By R . O. King , M .A .Sc, and H. Moss, D.Sc.Communicated by the Director of Scientific ResearchA ir Ministry. R . * M . No. 1362 (E . 44). (11 pagesand 8 diagrams.) July, 1930. Price 9d . net.It was shown by experiments described earlier* that the high anti-knockvalue given to fuels by the addition of benzole or metallic dope was generallydiminished when lubricating oil was distributed throughout the fuel-air

mixture d uring combustion. The several oils used in these experimentswere, the vege table oils, rap e an d cas tor, oleine and oleic acid, also fatty nils,but usually derived from animal sources, and mineral oils as represented hytwo proprietary blends. The detonation inducing actions of the three typesof oil differed remarkably, and varied with the substance used to increase theanti-knock property of the original fuel.The present experimental work was, therefore, undertaken (o) to ascertainthe effect on detonation of typical basic varieties of mineral oils, and (b ) toinvestigate the difference in oil effect on plain and doped fuels.The typical oils used we re: R ussian No . 1 as a naphthenio base oil, anoil from a Venezuela crude as representing an asphaltic base oil , a lightand a heavy distillate from a Pennsylvania paraffin base, and a refinedresidual oil from the same cru de. The fuels were aviation petrol plusbenzole or lead dope, and a series of blends made up with varying proportionsof paraffins, naphthenes an d aromatics. The experiments covered numerouscombinations of the various oils and fuels, and were made with inductiontempera tures ranging from normal to 90 C.Considering fuels doped with ethyl fluid and used at normal inductiontemperature, the various oils differ in deleterious action, the asphaltic baseoil having the greatest and the paraffin residual (cylinder stock) the leasteffect. When th e induction tem per atu re Is raised all of the oils becomeequally deleterious in effect. With dope d fuels In gener al a given quan tityof oil tends to neutralise a definite equivalent quantity of dope as the com-pression ratio is Increased. Wh en th e oils are used with undoped fuels thedeleterious oil effect increases with the compression ratio at which the fuelsare used, within a range of 0 10 ratio, and within this range fuels containingnaphthenes in large proportion are less deleteriously affected than fuelscontaining equivalent propo rtions of benzole.Incidentally, i t was found tha t cyclohexane, a nap hthenic fuel, was lessdeleteriously affected by high Induction temperature than other undopedfuels given an equal initial anti-knock value by add itions of benzole. Further,when eyclohexane was used as a fuel, maximum detonation was obtained atweak mixtures, a cha racteristic previously thou ght to be peculiar to dopedfuels.

King and Moss, R . & M. 1318, " Detonation and L ubricating Oil."M A X I M U M L I F T I N C L O S E D A N D O P E N J E T T U N N E L S .By F. B. Bradfield, Math, and Nat. Sci. Triposes,K. W. Clark, B.Sc., and R . A . Fairth orne. Communi-cated by the Director of Scientific Research, A ir

Ministry. R . & M. No. 1363 (A e. 491). (19 pagesand 6 diagrams.) December, 1930. Price Is . net.The three pa rts of this repor t record various tests in connection with tunne 1Interference on the m aximum lift of aerofoils. No general rule has beenfound governing this effect. The results may be summ arised thus :(o) The ma ximu m lift coefficient of a slotted wing ma y be Increasedconsiderably by tunne l Interference. Fo r a 6 x 36 in. wing, i L max.,measured in a 4-ft . tunnel, may be greater by 0-1 or more, than whenmeasured Jn a 7-ft. tun nel.(6) The maximum lift coefficient of a wing of lower lift coefficient, such anB.A .F.30, B..A.F.82, Aerofoil A (a modified form of R .A.F.15) increases withdecreasing sUe of tunne l, but to a much lesser extent. Taking the samecomparisons as before, the increase wonld be 0 025, or a quarter of that foundfor the slotted wing.(e) The maximum lift coefficient measured in an open jet tunnel varieslittle from the free air value.l'Xow o r A IR ADJACENT TO THE SURFA CE o r A R O T A T I N G

C Y L I N D E R . By E. G. R i c h a r d s o n , B. A . , Ph . D . , D . Sc .R . & M . No." 1368 (A e. 495). (12 pages and 18 dia-grams.) December, 1930. Price Is . ne t .Measurements of the average velocity and the amplitude of velocityfluctuation close to the surface of a rotating cylinder in a stream in two-dimensional flow at various ratios of stream to peripheral velocity are made,using the methods of E. & M. 1224.*It is shown that the results are in agreement with the classical theory ofthe Magnus effect outside the boundary layer, provided a circulation equalto abou t two-thirds of the theoretical value is assumed. These results areco-ordinated with measurements of lift and drag of the rotating cylinder ona force balance. The variation of skin friction roun d the surface is com-pared with the variation of the power required to rotate the cylinder.The general results are in agreement with a modified form of the circulationhypothesis, having regard to the steep gradient of velocity in the boundarylayer round the cylinder.

t 60-9 Incidence. It should be noted, however, that the m odel was notfitted with slots. R . & M. 1224. " On the Flow of Air A djacent to the Surface of anAerofoil," by N. A. V. Piercy and F. G. Richardson.726 h