POF • • •
• •
•
LFnew/LForiginal = (Vnew/Vo (Vnew/Voriginal)^2 Flap extension causes a reduction in stall speed and the maximum glide distance. he Principle o! "ontinuit# states$ %&! the cross sectional area o! a streamlined !low o! su'sonic air is increased the !low elocit# will decrease%. he "ontinuit# e*uation is$ %"ross+sectional area (,) x Velocit# Velocit# (V) = "onstant%. he illustration shows the relationships 'etween ",- (") ,,- () and ach num'er () with changing Pressure ,ltitude. &t can 'e seen that when in a clim' a'oe the tropopause at a constant ach num'er the ,- must remain constant.
•
V- is the stall speed in the landing con!iguration
•
V-0 is the stall speed in a speci!ied con!iguration
•
V-0g is the minimum speed at which the aeroplane can deelop a li!t !orce (normal to the !light path) pa th) e*ual to its 1eight 1eight
•
he re!erence stall speed is V-
•
he &nduced 3rag coe!!icient !ormula is$ "3i = ("L)2 / ,spect atio
•
ip ortices are wea4er when the aircra!t is close to the ground (within hal! a wing span). he illustration shows that when an aeroplane is ou t o! ground e!!ect the induced downwash (angle) is increased and conse*uentl# the induced angle o! attac4 is increased.
•
•
•
•
he angle o! attac4 o! an aero!oil section is de!ined as the angle 'etween the undistur'ed air!low and the chord line. ,s !laps are deplo#ed the wing cam'er increases and sometimes in addition the wing area. -o i! a constant angle o! attac4 is maintained the Li!t coe!!icient would increase. he illustration shows a slat (a moea'le part o! the leading edge which when actiated !orms a slot at the wing leading edge). e dge). he slot thus !ormed directs high energ# air onto the wing upper sur!ace which increases the 'oundar# la#er 4inetic energ# on the top o! the wing and dela#s the stall to a higher angle o! attac4 &t can 'e seen !rom the illustration that slats do not change the cam'er o! the wing 'ecause the Li!t cure is merel# extended
egarding su'sonic air!low in a entur#$ 0. the d#namic pressure in the undistur'ed !low and in the throat are not e*ual
• •
2. he total pressure in the undistur'ed !low and in the throat are e*ual 5. -tatic pressure acts in all directions -tall speed (&,-) aries with weight. he Li!t !ormula is$ L = 0/2rho V2 "L - where rho is the air densit# V is the rue ,irspeed ,irspeed (,-) "L is the Li!t coe!!icient and - is the wing area. ogether ogether 0/2rho V2 = 3#namic Pressure (*). ,lthough the *uestion as4s #ou to consider the Li!t !ormula it is actuall# necessar# to consider how the ,irspeed &ndicator wor4s. 1e 4now !rom the 3#namic Pressure !ormula (* = 0/2rho V2) i! we dou'le the speed o! the aircra!t through the air (dou'le the ,-) 3#namic pressure (*) will 'e !our times times greater 'ut 'ecause o! the s*uare root gearing inside the ,-& the &ndicated ,irspeed ,irspeed will onl# dou'le. &! #ou now put !our times the 3#namic Pressure into the Li!t !ormula it is clear that the Li!t will 'e !our times greater. &! #ou pre!er the math6s explanation &,- is proportional to the s*uare root o! 3#namic Pressure (*). ransposed this 'ecomes 3#namic P ressure (*) is proportional to &,- s*uared. his is a theor# *uestion. &n !light #ou nee d to 4eep Li!t e*ual to 1eight. -o to maintain constant Li!t as airspeed (&,-) is dou'led #ou would need to reduce the angle o! attac4 to 0/7 o! o ! its preious alue.
•
&nduced 3rag is caused '# tip ortices. he stronger or more e!!ectie the tip ortices the greater the &nduced 3rag. ip ortices !orm 'ecause o! the pressure di!!erential 'etween the top and 'ottom sur!ace o! the aero!oil. he greater the 1eight 1e ight the greater the Li!t re*uired to 'alance it there!ore the top and 'ottom sur!ace pressure di!!erential will 'e greater giing stronger tip ortices and more &nduced 3rag. "onse*uentl# i! wing Li!t is 8ero there will 'e no tip ortices and there!ore 8ero &nduced 3rag.
•
he relationship 'etween pressure densit# and a'solute temperature o! a gien mass o! air can 'e expressed as p / (rho 9 ) = constant. his *uestion relates to the :&deal ;as Law: which states that air densit# (rho) is proportional to pressure (p) and inersel# proportional to a'solute temperature (). ,ssuming no compressi'ilit# e!!ects induced drag at constant &,- is a!!ected '# ,eroplane 1eight 1eight (mass) will a!!ect the amount o! Li!t produced so will there!ore a!!ect &nduced 3rag. 1here the hori8ontal axis crosses the ertical axis is 8ero Li!t coe!!icient ("L). he swept 'ac4 wing has an increased tendenc# to stall !irst at the tips due to the span+wise !low !rom root to tip. his reduces Li!t a!t o! the "; and ge nerates an aircra!t nose up pitching moment. he nose+up pitching moment o! some aircra!t with a swept 'ac4 wing is so iolent and !ast that no human is capa'le o! reacting !ast enough to preent it. his is 4nown as a 3eep -tall which can also 'e called a -uper -tall. ,n# aeroplane suscepti'le to 3eep -tall - tall can neer 'e allowed to stall and must there!ore 'e !itted with a %-tall preention deice% called a -tic4 Pusher. 1ith 1ith
•
• •
•
•
•
•
increasing angle o! attac4 the -tall 1arning s#stem (-tic4 sha4er) will actiate as normal 'ut i! the pilot does not decrease the angle o! attac4 the stic4 pusher then actiates and pushes the stic4 (eleator control) !orward to preent 3eep -tall !rom happening. &n two dimensional !lows span wise !low is not considered< there!ore there ca n 'e no &nduced 3rag. &nter!erence 3rag results !rom the inter!erence o! the 'oundar# la#ers o! adacent components 'ut here we hae no adacent components and there!ore we don:t need to consider &nter!erence 3rag. he onl# correct answer is Pressure 3rag and -4in Friction 3rag. Flap as#mmetr# usuall# happens when high li!t deices are selected + one side moes and the other side sta# in its original position. he illustration shows that !lap as#mmetr# will cause a large rolling moment at an# angle o! attac4 whereas slat as#mmetr# would merel# cause a large di!!erence in "L,>. One o! the 'iggest adantage o! a slat is the# do not signi!icantl# increase 3rag. he gien :correct: answer there!ore ma4es no sense. %Flap as#mmetr# causes a large rolling moment% is a'solutel# correct 'ut not$ %a #awing moment !rom slat as#mmetr#%.
•
From the illustration it can 'e seen that L/3 ,> corresponds to the speed V3 in straight and leel (0g) !light. -o i! speed is reduc ed 'elow L/3 ,> the otal 3rag will increase due to increasing &nduced 3rag.
•
he stall speed decreases$ (all other releant !actors are constant) when during a manoeure the aeroplane nose is suddenl# pushed !irml# downwards (e.g. as in a push oer). Load Factor (n) = Li!t / 1eight. &! the aircra!t is manoeured '# suddenl# pushing the nose down the Li!t will 'e reduced and !rom the a'oe !ormula it can 'e seen that less Li!t will decrease the stall speed. ?ecause tip ortices generate &nduced 3rag an#thing that inter!eres with the !ormation or strength o! tip ortices will reduce &nduced 3rag. 1inglets reduce the strength o! tip ortices and there!ore reduce &nduced 3rag. @oweer an#thing placed in the air!low will also generate Parasite 3rag.
•
•
@igh li!t deices (leading and trailing edge) are !itted to an aeroplane to decrease the ta4e+o!! and landing distance. 3ecreasing the ta4e+o!! and landing distance is achieed '# increasing the maximum li!t coe!!icient ("L,>) which will decrease the stall speed and there!ore the minimum operating speed. 1hen (trailing edge) !laps are deplo#ed !rom A to 0BA there will 'e a comparatiel# large increase in "L and a small increase in "3 'ut with each successie increase in !lap angle the increase in "L will 'e less whereas the increase in "3 will 'e greater. For maximum e!!icienc# during ta4e+o!! not onl# is an increase in "L,> desired 'ut an# increase in 3rag needs to 'e minimi8ed. "onse*uentl# the :optimum: !lap !or ta4e+o!! is approximatel# 0BA. @oweer !or landing maximum 3rag is re*uired so !ull !lap is necessar#.
•
•
•
Cse the Li!t !ormula !or this *uestion$ (L = 0/2rho V2 "L -) where - is wing area. &! the wing area is increased Li!t will increase 'e cause it is directl# proportional to wing area. -weeping the wing either !orward or as is more usual 'ac4wards decreases the aerod#namic e!!icienc# o! the wing. "onse*uentl# "L,> is decreased. here!ore i! sweep angle is increased "L,> will decrease. &! we consider V- at "L ,> (V- = the s*uare root o! L / 0/2rho "L ,> -) it can 'e seen that a decrease in "L,> will increase the stall speed.
&ncreasing !orward sweep increases the stall speed. 1ing anhedral is used '# the designers to gie an aeroplane the re*uired leel o! Lateral -tatic -ta'ilit#. ,nhedral has no in!luence on stalling. • •
•
•
he load !actor is greater than 0 (one) when li!t is greater than weight. 1ing !ences reduce the span+wise !low and help to reduce the increased tendenc# !or a swept wing to tip stall and conse*uentl# pitch+up is reduced. 1ing !ences there!ore improe the low speed handling characteristics o! a swept wing. Cse the Li!t !ormula to consider this *uestion$ (L = 0/2rho V2 "L -) where :rho: is air densit#. ,ltitude is the onl# aria'le in the *uestion and an # change in altitude will a!!ect the air densit#. he lower the altitude the greater the air densit# which will re*uire a lower ,- to maintain a constant &,- and Li!t e*ual 1eight. he swept 'ac4 wing has an increased tendenc# to stall !irst at the tips due to the span+wise !low !rom root to tip. his reduces Li!t a!t o! the "; and ge nerates an aircra!t nose up pitching moment. he nose+up pitching moment o! some aircra!t with a swept 'ac4 wing is so iolent and !ast that no human is capa'le o! reacting !ast enough to preent it. his is 4nown as a 3eep -tall which can also 'e called a -uper -tall. ,n# aeroplane suscepti'le to 3eep -tall can neer 'e allowed to stall and must there!ore 'e !itted with a %-tall preention deice% called a -tic4 Pusher. 1ith increasing angle o! attac4 the -tall 1arning s#stem (-tic4 sha4er) will actiate as normal 'ut i! the pilot does not decrease the angle o! attac4 the stic4 pusher then actiates and pushes the stic4 !orward to preent 3eep -tall !rom happening. , wing with !orward sweep does not hae an increased tendenc# to tip stall and will not generate a nose+up pitching moment 'e!ore the stall. here!ore an aeroplane with a swept !orward wing is not suscepti'le to 3eep -tall. , contri'utor# !actor to 3eep -tall is a +tail. ,t the extremel# high angle o! attac4 su!!ered in a 3eep -tall the +tail would 'e immersed in separated air!low ma4ing the eleator ine!!ectie and preenting the pilot !rom decreasing the angle o! attac4. , low mounted tail would not su!!er !rom this pro'lem and is there!ore not a contri'utor# !actor to 3eep -tall.
•
From the illustration it can 'e seen that as the angle o! attac4 decreases the stagnation point moes up on the leading edge and the point o! minimum pressure moes a!t.
•
ip ortices induce downwash 'ehind the wing at the tip. his induced downwash changes the aerage direction o! the air!low oer the wing. he angle o! this modi!ied air!low (he D!!ectie ,ir!low) to the elatie ,ir!low is the &nduced ,ngle o! ,ttac4. -ee illustration 'elow.
•
,ssuming constant &,- when an aeroplane enters ground e!!ect the e!!ectie angle o! attac4 increases. "L = 0 / VE ,ir densit# is mass per unit olume< the unit !or mass is the 4g (4ilogram) and the unit !or olume is the m (cu'ic meter) hence densit# is 4g/m.
• •
he unit !or !orce is the Gewton (G). •
1ing loading is the aircra!t weight (G) diided '# wing area (mE) hence wing loading is G/mE. 3#namic pressure is !orce per unit area so is also G/mE.
• •
Li!t is generated when a certain mass o! air is accelerated in its !low direction . he d#namic pressure increases as static pressure decreases.
•
he tur'ulent 'oundar# la#er has more 4inetic energ# than the laminar 'oundar# la#er.
•
, positie cam'er aero!oil section at 8ero angle o! attac4 will generate a small amount o! li!t. his is 'ecause the cross sectional area o! the streamlined !low is accelerated more oer the :top: sur!ace than the 'ottom. From the illustration it can 'e seen that the Li!t cure o! a positie cam'ered aero!oil intersects the ertical axis o! the "L + alpha graph a'oe the point o! origin.
•
On a swept wing aeroplane at low airspeed the %pitch up% phenomenon is caused '# wingtip stall.
•
•
•
•
•
•
•
• •
•
•
• • •
he angle 'etween the aeroplane longitudinal axis and the chord line is the angle o! incidence. From the illustration 'elow it can 'e seen that the aeroplane (otal) 3rag in straight and leel !light is lowest when the Parasite 3rag is e*ual to the &nduced 3rag. 3ue to the !ormation o! shoc4 waes on the wing top and 'ottom sur!ace and the su'se*uent rearwards moement o! the shoc4 waes as the aeroplane accelerates !rom su'sonic to supersonic speed the distri'ution o! pressure on the wing changes causing 'oth the centre o! pressure ("P) and the aerod#namic centre (,") to moe a!t !rom 2BH to the BH chord position (mid chord). , s#mmetrical aero!oil section at 8ero degrees angle o! attac4 will accelerate the air oer the top sur!ace and the 'ottom sur!ace '# the same amount hence there will 'e no net li!t either upwards or downwards co nse*uentl# the li!t coe!!icient will 'e 8ero. ,n aeroplane maintains straight and leel !light while the &,- is dou'led. he change in li!t coe!!icient will 'e x .2B. he most important pro'lem o! ice accretion on a transport aeroplane during !light is reduction in "Lmax. he e!!ect o! hea# (tropical) rain on the aeroplane is signi!icant. he large amount o! water on the aeroplane will increase the 1eight and the accumulation o! water will also disrupt the air!low. his will 'oth decrease "L ,> and increase 3rag. ,t the highest alue o! the li!t/drag ratio the total drag is lowest. ?# de!inition li!t is perpendicular to the relatie air!low (also 4nown as relatie wind or !ree stream !low) and drag is parallel to an d in the same direction as the relatie air!low (relatie wind or !ree stream !low). he Parasite 3rag !ormula is$ 3p = 0/2rho V2 "3p - there!ore 3p is proportional to speed s*uared. ,n aeroplane accelerates !rom I 4t to 0J 4t at a load !actor e*ual to 0. he induced drag coe!!icient (i) and the induced drag (ii) alter with the !ollowing !actors$ (i) 0/0J (ii) 0/7 &nduced drag is created '# the spanwise !low pattern resulting in the tip ortices. Vortex generators trans!er energ# !rom the !ree air!low into the 'oundar# la#er . ,nhedral can 'e used '# the designers to set the re*uired Lateral -tatic - ta'ilit#. &ncreasing ,nhedral reduces Lateral -tatic -ta'ilit# 'ut has no in!luence on stall speed. &ncreasing the sweep angle o! the wing is used '# the designers to increase the "ritical ach num'er. &ncreasing the sweep angle also ma4es the wing less aerod#namicall# e!!icient which will increase the stall speed. , +tail is !itted '# the designers to reduce the in!luence o! wing downwash on the hori8ontal sta'ili8er. , +tail also acts as an %Dnd Plate% on the !in which ma4es the !in more aerod#namicall# e!!icient. , +tail has no in!luence on the stall speed.
•
,n %erect% spin is with the top o! the aeroplane towards the inside o! the helix (cor4screw shape) o! the spin. he %-tandard% spin recoer# includes$ Full rudder opposite to the direction o! rotation close the throttle neutrali8e the roll control and moe the pitch control !orward. -o 4eeping the aileron control neutral during recoer# is the onl# true statement here.
•
•
•
•
%Kust 'e!ore the stall a nose down pitching moment is generated%. his is caused '# the rearward moement o! the "P. he ane o! a stall warning s#stem with a !lapper switch is actiated '# the change o! the stagnation point. &n a stead# leel co+ordinated turn the Li!t must 'e greater than the 1eight so there!ore the Load Factor (n) will 'e greater than 0. o consider the e!!ect on the stall speed we can re!er to the !ormula$ V- GD1 = V- OL3 x the s*uare root o! the Load Factor. he 4e# to this *uestion is the part which states % whilst maintaining leel !light at a constant &,- % o maintain leel !light at a constant &,- Li!t must remain the same 'ut when !laps are extended Li!t increases. "onse*uentl# in order to maintain leel !light as the !laps go down the angle o! attac4 must 'e decreased to 4eep Li!t constant. here!ore when !laps are extended whilst maintaining straight and leel !light at constant &,- the Li!t coe!!icient ("L) must remain the same.
•
•
railing edge !lap extension will decrease the critical angle o! attac4 an d increase the alue o! "Lmax. &t can 'e seen !rom the illustration that deplo#ing a slat will !orm a slot and deplo#ing a Mrueger !lap does not !orm a slot. &t is also eident that deplo#ing 'oth a Mrueger !lap and a slat will increase the critical angle o! attac4.
•
&t can 'e seen !rom the illustration that o! the t#pes o! trailing edge !lap listed the Fowler !lap is the most e!!ectie (highest "L,>).
•
, slotted !lap will increase the "Lmax '# increasing the cam'er o! the aero!oil and re+energi8ing the air!low.
•
he purpose o! an auto+slat s#stem is to extend automaticall# when a certain alue o! angle o! attac4 is exceeded. he t#pe o! stall that has the largest associated angle o! attac4 is a deep stall. ,s an aeroplane accelerates through the transonic region the shoc4 waes on the wing moe rearwards and change the pressure distri'ution on the wing which ma4es the ,erod#namic "entre (,") moe rearwards !rom 2BH a!t towards BH a!t.
• •
he top illustration shows the Longitudinal -tatic -ta'ilit# in su'sonic !low with the ," at 2BH a!t. he 'ottom illustration shows the ," at BH a!t when the !low 'ecomes supersonic.
&t can 'e seen that in supersonic !low when the ," has moed a!t to the BH chord position the desta'ili8ing wing moment is reduced and conse*uentl# -tatic Longitudinal -ta'ilit# is increased.
•
•
•
•
•
•
he shoc4waes associated with supersonic !light cause the p ressure patterns to appear rectangular in !orm unli4e the pressure patterns !or su'sonic !light which appear smooth and slightl# humped when plotted on a graph. &ncreasing air densit# will hae the !ollowing e!!ect on the drag o! a 'od# in an air stream (angle o! attac4 and ,- are constant) the drag increases. %, line connecting the leading+ and trailing edge midwa# 'etween the upper and lower sur!ace o! a aero!oil%. his de!inition is applica'le !or the cam'er line. Power is the rate o! doing wor4 (how *uic4l# wor4 is done). 1or4 is Force (G) x 3istance (metre) so Power is Force (G) x 3istance (metre) diided '# time (seconds) or G. m/s "ompared with the clean con!iguration the angle o! attac4 at "Lmax with trailing edge !laps extended is smaller. he illustration shows an Dxpansion 1ae. One o! the characteristics o! supersonic !low is that it can !ollow a conex corner 'ecause it expands upon reaching the corner. "onse*uentl# elocit# increases and the other parameters pressure densit# and temperature all decrease. Local speed o! sound (a) is proportional to air temperature. -o as the temperature decreases through an expansion wae the local speed o! sound will decrease.
• •
• • •
•
•
•
•
•
he !low on the upper sur!ace o! the wing has a component in wing root direction. -poiler extension increases the stall speed the minimum rate o! descent and the minimum angle o! descent. 3uring a clim'ing turn to the right the angle o! attac4 o! 'oth wings is the same. Pstat N rhoVE= constant. ;ien an initial condition in straight and leel !light with a speed o! 0.7 V-. he maximum 'an4 angle attaina'le without stalling in a stead# co+ordinated turn whilst maintaining speed and altitude is approximatel#$ JA. 3uring a stead# hori8ontal turn the stall speed increases with the s*uare root o! the load !actor. he illustration shows the correct orders o! increasing critical angle o! attac4 a re$ !laps onl# extended clean wing slats onl# extended.
Floating due to ground e!!ect during an approach to land will occur when the height is less than hale o! the length o! the wing span a'oe the sur!ace. ,n aero!oil is cam'ered when the line which connects the centres o! all inscri'ed circles is cured. From the illustration it can 'e seen that while !l#ing in the speed unsta'le region an# reduction in speed will gie an increase in 3rag which will lead to a greater reduction in speed.
•
•
•
•
,n aeroplane in straight and leel !light is su'ected to a strong ertical gust. he point on the wing where the instantaneous ariation in wing li!t e!!ectiel# acts is 4nown as the$ aerodynamic centre of the wing. he ?oundar# La#er is a la#er o! air !lowing oer a sur!ace the lowest la#er o! molecules haing 8ero elocit# relatie the sur!ace 'ut increasing in speed with increasing distance !rom the sur!ace until the !ree stream !low elocit# is reached. here!ore the ?oundar# La#er has a lower aerage elocit# than that o! the !ree stream. he polar cure o! an aero!oil section is a graphic relationship 'etween li!t coe!!icient "l and drag coe!!icient "d. he ?oundar# La#er is a la#er o! air !lowing oer a sur!ace the lowest la#er o! molecules haing 8ero elocit# relatie the sur!ace 'ut increasing in speed with increasing distance !rom the sur!ace until the !ree stream !low elocit# is reached. here are two t#pes o! 'oundar# la#er the laminar 'oundar# la#er and the tur'ulent 'oundar# la#er. ,t the leading edge the 'oundar# la#er is initiall# laminar (no intermixing o! adacent la#ers) and will remain laminar as it moes rearwards until it 'ecomes tur'ulent at the ransition Point. he tur'ulent !low will continue towards the trailing edge and ust 'e!ore it reaches the trailing edge it will start to separate !rom the sur!ace this is called he -eparation Point. ?ecause the laminar 'oundar# la#er is smooth with no intermixing its 4inetic energ# is comparatiel# low and is there!ore easil# separated. ?ecause the tur'ulent 'oundar# la#er has lots o! intermixing 'etween the la#ers o! air molecules it will there!ore contains more 4inetic energ# than a laminar 'oundar# la#er. "onse*uentl# compared to a laminar 'oundar# la#er a tur'ulent 'oundar# la#er is 'etter a'le to resist a positie pressure gradient 'e!ore it separates.
•
•
• • •
, positiel# cam'ered aero!oil will generate 8ero li!t$ at a negatie angle o! attac4. 1ing Loading = 1eight / 1ing ,rea. here!ore !or a gien 1ing ,rea the greater the 1eight the higher the 1ing Loading. -tall speeds are determined with the "; at the !orward limit. inimum control speeds are determined with the "; at the a!t limit. egulations state that the stall warning must 'egin at a speed (V-1) exceeding the speed at which the stall is identi!ied '# not less than B 4t or BH ",- whicheer is greater. hat would 'e 0.B V- so the onl# correct answer is %greater than V-%.
•
•
•
he ean ,erod#namic "hord (,") !or a gien wing o! an# plan !orm is the chord o! a rectangular wing with same moment and li!t. ?ehind the transition point in a 'oundar# la#er the mean speed and !riction drag increases. he !ormula we need when calculating Li!t (L) or Load Factor (n) in a turn is$ L or n = 0 / "O- P@&< where P@& = 'an4 angle. ,nd when calculating the a!!ect o! 'an4 angle on stall speed the !ormula 'ecomes V- GD1 = V- OL3 x the s*uare root o! 0 / "O- P@& or to put it another wa# V- GD1 = V- OL3 x the s*uare root o! n. 1here P@& is the 'an4 angle and n = Load Factor. here!ore when a pilot ma4es a turn in hori8ontal !light the stall speed increases with the s*uare root o! the Load Factor.
•
•
• • •
•
he true airspeed (,-) is lower than the indicated airspeed (&,-) at &-, conditions and altitudes 'elow sea leel. he li!t !orce acting on an aero!oil$ (no !low separation) is mainl# caused '# suction on the upper side o! the aero!oil. he relatie thic4ness o! an aero!oil is expressed in H chord. ,spect ratio o! a wing is the ratio 'etween wing span s*uared and wing area. he wing o! an aeroplane will neer stall at low su'sonic speeds as long as.... the angle o! attac4 is smaller than the alue at which the stall occurs.
he illustration shows an Dxpansion 1ae. One o! the characteristics o! supersonic !low is that it can !ollow a conex corner 'ecause it expands upon reaching the corner. "onse*uentl# elocit# increases and the other parameters pressure densit# and temperature all decrease. "onse*uentl# the densit# in !ront o! an expansion wae is higher than 'ehind it and the pressure in !ront o! an expansion wae is also higher than 'ehind it.
•
-tatic sta'ilit# is the initial (split second) reaction o! the aeroplane immediatel# !ollowing the remoal o! a distur'ing !orce. 3#namic sta'ilit# is what happens a!ter the initial reaction.
&! an aeroplane is staticall# sta'le it will start 'ac4 towards its preious state o! e*uili'rium and i! an aeroplane is staticall# unsta'le it will start to moe !urther awa# !rom its preious state o! e*uili'rium. here!ore 3#namic sta'ilit# is possi'le onl# when the aeroplane is staticall# sta'le. •
&t must 'e !ull# understood that this is a theor# *ue stion. &n practice #ou would not een taxi with ice on the wing leading edge let alone attempt a ta4e+o!!. ,'out IH o! Li!t is generated '# the top sur!ace o! the wing. ,nd o! the top wing sur!ace a'out 2H a!t o! the leading edge is where the maorit# o! the !low acceleration ta4es place. -o an#thing which inter!eres with the air!low oer the leading 2H o! the wing top sur!ace will hae a maor in!luence on Li!t generation. he *uestion gies us ice located on the wing leading edge. 3uring the ta4e+o!! run the aeroplane is merel# accelerating do wn the runwa# and ice located on the wing leading edge will onl# increase the 3rag slightl#. 3uring the clim' with all engines operating the aeroplane will 'e at a small angle o! attac4 (a'out 7A) and ice on the wing leading edge will increase the 3rag and decrease "L '# a moderate amount. 3uring the last part o! rotation the angle o! attac4 will 'e *uite high and an# ice on the leading edge will hae a signi!icant negatie in!luence on Li!t production and in !act the wing will pro'a'l# !ull# stall.
•
he !ollowing !actors increase stall speed$ an increase in load !actor a !orward cg shi!t and decrease in thrust.
•
he illustration shows a (Gormal) shoc4 wae. ,s the air!low passes through a shoc4 wae the elocit# decreases and the pressure the densit# and the temperature all increase.
•
,t what speed does the !ront o! a shoc4 wae moe across the earth:s sur!ace he ground speed o! the aeroplane. Low speed pitch up is caused '# the outward dri!t o! the 'oundar# la#er on a swept+'ac4 wing. he "entre o! Pressure ("P) is the point on an aero!oil through which the Li!t acts. 1e 4now that the location o! the "P is a product o! the aerage o! the top
•
•
sur!ace Li!t pressure and the 'ottom sur!ace Li!t pressure. 1e are also aware that as the angle o! attac4 is increased the Li!t pressure distri'ution changes and the "P moes !orward. 1hen stud#ing Longitudinal -tatic -ta'ilit# and certain aspects o! ,ircra!t Limitations use o! the moing "P would 'e con!using so in those circumstances a di!!erent re!erence point is used called the ,erod#namic "entre (,"). "onsider illustration (,) 'elow$ he ,erod#namic "entre (,") is located 2BH a!t Qo! the leading edgeR. &t can 'e seen that at angle o! attac4 (alpha 0) Li!t (L0) acts through distance (d0) !rom the ,erod#namic "entre (,"). his generates a nose down pitching moment () a'out the ,". &! #ou now consider illustration (?)$ it can 'e seen that an upward ertical gust has increased the angle o! attac4 to (alpha 2) which increases the Li!t to (L2) and moed the "P !orward decreasing the distance o! the "P !rom the ," to (d2). ?ut 'ecause (L2) has increased '# the same amount that (d2) has decreased the nose down pitching moment () a'out the ," has remained the same. Gow consider illustration (")$ 'ecause pitching moment () has sta#ed the same with an increase in angle o! attac4 the change in Li!t (delta L) can 'e considered to act at the ," as illustrated in illustration (3). Csing the ,erod#namic "entre (,") as the point through which the change in Li!t acts greatl# simpli!ies the isuali8ation o! the e!!ect o! changes in angle o ! attac4 due to gusts 'oth in the stud# o! Longitudinal -tatic -ta'ilit# and the a!!ect o! a gust stressing the aeroplane when stud#ing ,ircra!t Limitations. Facts to remem'er a'out the ,erod#namic "entre$ 0 he ," is located 2BH a!t Qo! the leading edgeR in su'sonic !low. 2. he pitching moment a'out the ," is alwa#s nose down. 5. he pitching moment a'out the ," does not change i! the angle o! attac4 changes. 7. he ," is the point through which the change in Li!t acts Qdue to a change in angle o! attac4R.
•
1ing sweep angle is the angle 'etween the *uarter+chord line o! the wing and the lateral axis.
•
he mean geometric chord o! a wing is the wing area diided '# the wing span. aper ratio o! a wing is the ratio 'etween tip chord and root chord. 1ing twist (geometric and aerod#namic) is used to$ 0. improe stall characteristics. 2. educe induced drag. ?ernoulli:s e*uation is Pstat N 0/2 9 rho 9,-E = constant. &! ice is present on the leading edge o! the wings it ma# increase the landing distance due to a higher Vth with 7+BH. he load !actor is less than 0 (one) during a stead# wings leel descent. he load !actor is less than 0 (one) 1hen li!t is less than weight. he load !actor is less than 0 (one) during a stead# wings leel clim'. he top shoc4 will 'e at its wea4est at crit it is onl# ust !orming he shoc4wae moes a!t as increases ,t 0 the shoc4wae has moed to the trailing edge 1hen the aileron is de!lected down the air!low oer the top increases and the shoc4wae moes a!t in the increased !low ,s altitude increased the stall speed (&,-) initiall# remains constant and at higher altitudes increases.
• •
• •
• • • • • • •
•
•
• •
-hoc4 induced separation can occur 'ehind a strong normal shoc4 wae independent o! angle o! attac4. he su'sonic speed range ends at crit. here are three aeroplane speed ranges. he slowest speed range is :-u'sonic:< de!ined as the speed range when all o! the !low speeds relatie to the sur!ace o! the aeroplane are less than the speed o! sound (L S 0.). he :su'sonic: speed region is 'etween 8ero and the "ritical ach num'er ("&). he next highest speed range is :ransonic:< de!ined as the speed range where some o! the !low speeds relatie to the sur!ace o! the aeroplane are less than the speed o! sound (L S 0.) and some o! the !low speeds relatie to the sur!ace o! the aeroplane are greater than the speed o! sound (L T 0.). he :ransonic: speed region is 'etween the "ritical ach num'er ("&) and approximatel# 0.2 he highest speed region in which we are interested is :-upersonic:< de!ined as the speed range where all o! the !low speeds relatie to the sur!ace o! the aeroplane are greater than the speed o! sound (L T 0.). he :-upersonic: speed region is 'etween approximatel# 0.2 and B.
•
•
•
•
he critical ach num'er o! an aeroplane is the ach num'er a'oe which locall# supersonic !low exists somewhere oer the aeroplane. he critical ach num'er o! an aeroplane can 'e increased '# sweep'ac4 o! the wings. 1ae 3rag is mostl# caused '# the air!low 'eing heated as it passes through the shoc4 waes 'ut a small proportion o! 1ae 3rag is due to shoc4 wae induced separation. -wept 'ac4 wings are !itted to an aeroplane to increase its critical ach num'er ("&). here!ore decreasing wing sweep'ac4 will decrease "&. he 3rag 3iergence ach Gum'er is the ach num'er at which the aerod#namic 3rag on an aero!oil or an air!rame 'egins to increase rapidl# as the ach num'er continues to increase. his increase in 3rag is a result o! shoc4 wae !ormation (1ae 3rag) and can cause the 3rag coe!!icient to rise to more than ten times its low speed alue. he 3rag 3iergence ach Gum'er is usuall# close to and alwa#s greater than the "ritical ach num'er. here!ore i! the "ritical ach num'er decreases the 3rag 3iergence ach Gum'er will decrease.
•
,n o'li*ue shoc4 wae has similar a!!ects on the air!low passing through it as does a normal shoc4 wae 'ut there are a !ew small ariations. 0. ,n o'li*ue shoc4 wae is so called 'ecause it is at an angle o! greater than UA to the direction o! the upstream !low. 2. here &- a !low direction change through an o'li*ue shoc4 wae. 5. he air!low enters an o'li*ue shoc4 wa e at supersonic speed and emerges at a lower 'ut still supersonic speed.
7. here is an increase in -tatic Pressure 3ensit# and "ompression through an o'li*ue shoc4 wae. B. here is a decrease in the energ# o! the air!low and a decrease in otal Pressure through an o'li*ue shoc4 wae. J. he temperature o! the air increases through an o'li*ue shoc4 wae. For this *uestion we need to remem'er$ 'ecause the temperature increases through an o'li*ue shoc4 wae the temperature 'ehind is higher than in !ront o! it. ,nd 'ecause the -tatic Pressure increases through an o'li*ue shoc4 wae the -tatic Pressure 'ehind is higher than in !ront o! it. •
•
•
•
•
•
&. %uc4 under% is caused '# an a!t moement o! the centre o! pressure o! the wing. &&. %uc4 under% is caused '# a reduction in the downwash angle at the location o! the hori8ontal sta'ili8er. he illustration shows that in comparison to a conentional aero!oil section t#pical shape characteristics o! a supercritical aero!oil section are$ a larger nose radius !latter upper sur!ace and negatie as well as positie cam'er.
3uring a descent at a constant ach num'er (assume 8ero thrust and standard atmospheric conditions) the angle o! attac4 will decrease. &! a s#mmetrical aero!oil is accelerated !rom su'sonic to supersonic speed the aerod#namic centre will moe a!t to the mid chord. educing the thic4ness/chord ratio o! a wing increases "& in the same wa# as does a swept wing. here!ore there will 'e a reduction in the ariations in 3rag coe!!icient a dela# in the onset o! shoc4 wae !ormation and a reduction in the ariations in Li!t coe!!icient. , normal shoc4 wae has seeral distinguishing !eatures chie! among them 'eing$ 0. , Gormal shoc4 wae is so called 'ecause it is normal (perpendicular) to the direction o! the upstream !low. 2. here is no !low direction change through a normal shoc4 wae. 5. he air!low enters a normal shoc4 wae at supersonic speed and emerges at su'sonic speed. 7. here is a greater increase in -tatic Pressure 3ensit# and compression through a normal shoc4 wae compared to an o'li*ue shoc4 wae.
• •
•
•
•
•
•
•
•
• •
•
B. here is a greater decrease in the energ# o! the air!low and lower otal Pressure through a normal shoc4 wae compared to an o'li*ue shoc4 wae. J. he temperature o! the air increases through a normal shoc4 wae. . he least energ# loss through a normal shoc4 wae is when the Local ach num'er (L) is ust a'oe L0. "ompressi'ilit# e!!ects depend on ach num'er. he ach num'er is the ratio 'etween the ,- o! the aeroplane and speed o! sound o! the undistur'ed !low. , %-hoc4 -tall% is due to air!low separation !rom the !ormation o! a shoc4 wae. -hoc4 waes !orm at high speed and there!ore a small angle o! attac4. 1hen an aeroplane is !l#ing !aster than 0. an# pressure changes ta4ing place in the air !low oer the sur!ace will onl# a!!ect parts o! the aeroplane within the ach "one. ("&) is$ %he aircra!t ach num'er at which the local elocit# !irst reaches 0.%. he aerage modern high speed et transport has an "& o! a'out .I he loss o! total pressure in a shoc4 wae is due to the !act that 4inetic energ# in the !low is conerted into heat energ#. he maximum accepta'le cruising altitude is limited '# a minimum accepta'le load !actor 'ecause exceeding that altitude ur'ulence ma# induce mach 'u!!et. he speed o! sound aries with the s*uare root o! the a'solute temperature. he lower the temperature the lower the speed o! sound< and ice ersa. he :sonic 'oom: (or sonic 'ang) o! an aeroplane !l#ing at supersonic speed is a result o! the shoc4 waes the aeroplane is generating stri4ing the ear drums o! a listener (usuall# on the ground). ( = ,- / a). Vortex generators mounted on the upper wing sur!ace will decrease the shoc4 wae induced separation. ,s an aeroplane accelerates a'oe its critical ach num'er ("&) a shoc4 wae will !orm on the wing top sur!ace. ,s the aeroplane continues to accelerate this shoc4 wae will get thic4er increase in length and moe 'ac4wards. ,t a'out .U another shoc4 wae will !orm on the wing 'ottom sur!ace getting thic4er longer and moing 'ac4wards with !urther increases in ach num'er. ,t a'out .UI these two shoc4 waes will reach the trailing edge o! the wing. ,s soon as 0. is exceeded another shoc4 wae will !orm a small distance in !ront o! the wing. his is called the ?ow 1ae 'ecause it is similar in appearance to the 'ow wae in !ront o! a 'oat as it moes through water. 1ith !urther acceleration the ?ow 1ae will moe closer to the wing leading edge and eentuall# will get no closer at a'out 0.5 to 0.7. his is called$ the speed o! attachment.
•
• •
he sin o! the ach (cone) angle (W) = a / ,-. &n other words sin W = 0 / . here!ore the greater the ach num'er () the smaller the ach angle (W) he ach trim s#stem will adust the sta'ili8er depending on the ach num'er. 1hen air has passed through a shoc4 wae the local speed o! sound is increased.
•
• • •
•
• •
•
•
•
•
•
•
•
•
•
•
•
•
•
he speed range 'etween high and low speed 'u!!et increases during a descent at a constant &,-. @igh speed 'u!!et is induced '# 'oundar# la#er separation due to shoc4 waes. Vortex generators on the upper side o! the wing decrease wae drag. ach 'u!!et occurs at the ach num'er at which shoc4 wae induced 'oundar# la#er separation occurs. , transonic ach num'er is a ach num'er at which 'oth su'sonic and supersonic local speeds occur. -peed o! sound increases with temperature increase. &n transonic !light the ailerons will 'e less e!!ectie than in su'sonic !light 'ecause aileron de!lection onl# partl# a!!ects the pressure distri'ution around the wing. ,s an aircra!t accelerates through the transonic speed range the coe!!icient o! drag increases then decreases. -hoc4 stall occurs when the li!t coe!!icient as a !unction o! ach num'er reaches its maximum alue. 1hen an aeroplane is !l#ing !aster than 0. an# pressure changes ta4ing place in the air !low oer the sur!ace will onl# a!!ect parts o! the aeroplane within the ach "one. On a t#pical transonic air!oil the transonic rearward shi!t o! the "P occurs at a'out .IU to .UI. Go noticea'le shoc4 waes will !orm oer an# wing section when !l#ing ust a'oe "&. %"o!!in "orner% is the collo*uial name !or the pressure altitude where the speed !or low speed 'u!!et is the same as the speed !or high speed 'u!!et. he :echnical: name !or which is %he ,erod#namic "eiling%. 3utch oll is sta'ilit# related characteristic and occurs when the Lateral -tatic -ta'ilit# is greater than the 3irectional -tatic -ta'ilit#. -hoc4 waes hae no in!luence on 3utch oll. he local speed o! sound is dependent on temperature onl#. he lower the temperature the lower the local speed o! sound< and ice ersa. -tall speed does not ar# with 3ensit# ,ltitude. @oweer at er# high altitudes the stall speed will start to increase due to increasing ach num'er (compressi'ilit# e!!ects). he ach+trim !unction is installed on most commercial ets in order to minimi8e the aderse e!!ects o! changes in the position o! centre o! pressure. 1hat happens to lateral sta'ilit# when !laps are extended Lateral sta'ilit# is decreased. , downward adustment o! a trim ta' in the longitudinal control s#stem has the !ollowing e!!ect$ the stic4 position sta'ilit# remains constant. 3uring a phugoid altitude aries signi!icantl# 'ut during a short period oscillation it remains approximatel# constant.
•
•
• • •
,n increase o! 04t !rom the trimmed position at high speed has less e!!ect on the stic4 !orce than an increase o! 04t !rom the trimmed position at low speed. -tatic lateral sta'ilit# will 'e decreased '#$ the use o! a low rather than high wing mounting. 3utch roll will 'e corrected '# a #aw damper. -tatic lateral sta'ilit# will 'e decreased '#$ reducing wing sweep'ac4. his in!ormation proided is o! a "mcg diagram showing moment a'out the "; against alpha &n Part 0 the aircra!t wants to pitch up 'ut as alpha increases the pitch up !orce gets less and less i.e. it is sel! correcting and thus positiel# sta'le. 1here the trace crosses the hori8ontal line "mcg is 8ero the aircra!t neither wants to go nose up nor nose down. &n !act it is %in trim% in 'oth places. ,t Point 2 ust 'rie!l# changing alpha has no e!!ect on the nose up or down moment so the aircra!t is neutrall# longitudinall# sta'le. &n Part 5 un!ortunatel# !or the pilot an# increase in alpha will increase the nose up moment. he aircra!t is negatiel# sta'le + unsta'le. &n !act this is t#pical o! the entr# into an irrecoera'le deep stall. he slope o! the line shows sta'ilit# or not a'o e or 'elow the hori8ontal axis shows where the nose will want to go i! #ou let go o! the stic4
• • •
• •
•
-tatic lateral sta'ilit# will 'e increased '#$ increasing wing sweep'ac4. 1hat is predominantl# used to set the ail @ori8ontal -ta'ili8er !or a4e+o!! ";. -tatic directional sta'ilit# is the$ tendenc# o! an aeroplane to recoer !rom a s4id with the rudder !ree. Forward moement o! the "; will reduce control response and increase sta'ilit#. @ow can the designer o! an aeroplane with straight wings increase the static lateral sta'ilit# ?# increasing the aspect ratio o! the ertical sta'ili8er whilst maintaining a constant area. he pitching moment ersus angle o! attac4 line in the diagram which corresponds to a "; located at the neutral point o! a gien aeroplane at low and moderate angles o! attac4 is$ line 2.
he Geutral Point is the position o! the "; that gies the aircra!t neutral longitudinal static sta'ilit#.
• •
•
• •
•
• •
• • •
•
•
•
he e!!ect o! a high wing with 8ero dihedral is Positie dihedral e!!ect. he stic4 !orce per g o! a hea# transport aeroplane is 5 G/g. 1hat stic4 !orce is re*uired i! the aeroplane in the clean con!iguration is pulled to the limit manoeuring load !actor !rom a trimmed hori8ontal straight and stead# !light 7B G. 1hat is the recommended action !ollowing !ailure o! the #aw damper(s) o! a et aeroplane !l#ing at normal cruise altitude and speed prior to encountering 3utch roll pro'lems educe altitude and ach num'er. ,s the sta'ilit# o! an aeroplane increases its manoeura'ilit# decreases. Dxcessie static lateral sta'ilit# is an undesira'le characteristic !or a transport aeroplane 'ecause it would impose excessie demands on roll control during a sideslip. ,n aeroplane that tends to return to its pre+distur'ed e*uili'rium position a!ter the distur'ance has 'een remoed is said to hae positie static sta'ilit#. -tatic lateral sta'ilit# will 'e increased '# reducing wing anhe dral. he contri'ution o! the wing to the static longitudinal sta'ilit# o! an aeroplane depends on "; location relatie to the wing aerod#namic centre.
,n a!t "; shi!t decreases static longitudinal sta'ilit#. -tatic lateral sta'ilit# will 'e decreased '# increasing wing anhedral. For an aeroplane to possess d#namic sta'ilit# it needs static sta'ilit# and su!!icient damping. ,n aeroplane is sensitie to 3utch roll when static lateral sta'ilit# is much more pronounced than static directional sta'ilit#. he purpose o! a dorsal !in is to maintain static directional sta'ilit# at large sideslip angles. he e!!ect o! a positie wing sweep on static directional sta'ilit# is as !ollows sta'ili8ing e!!ect.
•
•
•
•
• •
•
•
•
• •
•
•
• •
•
• •
•
•
• •
&! the total sum o! moments a'out one o! its axes is not 8ero an aeroplane would experience an angular acceleration a'out that axis. ,n aeroplane has static directional sta'ilit#< in a side slip to the right initiall# the nose o! the aeroplane tends to moe to the right. For a normal sta'le aeroplane the centre o! grait# is located with a su!!icient minimum margin ahead o! the neutral point o! the aeroplane. he maximum a!t position o! the centre o! grait# is amongst others limited '# the re*uired minimum alue o! the stic4 !orce per g. he manoeura'ilit# o! an aeroplane is 'est when the cg is on the a!t cg limit. ,n aeroplane with an excessie static directional sta'ilit# in relation to its static lateral sta'ilit# will 'e prone to spiral die (spiral insta'ilit#). , ach trimmer corrects the change in stic4 !orce sta'ilit# o! a swept wing aeroplane a'oe a certain ach num'er. Positie static lateral sta'ilit# is the tendenc# o! an aeroplane to roll to the le!t in the case o! a sideslip (with the aeroplane nose pointing to the le!t o! the incoming !low). &! the static lateral sta'ilit# o! an aeroplane is increased whilst its static directional sta'ilit# remains constant its sensitiit# to 3utch roll increases. Lateral static sta'ilit# is determined '# ,ircra!t response to sideslip. he s#stem is concerned with positie :g: 'ut i! the pilot pushed the #o4e !orward and experienced negatie :g: it would increase the stic4 !orce the pilot !elt in pushing the #o4e !orward. &! an aeroplane exhi'its insu!!icient stic4 !orce per g this pro'lem can 'e resoled '# installing a 'o'weight in the control s#stem which pulls the stic4 !orwards. &n a stead# sideslip #ou are holding on aileron to 'alance the tendenc# o! the aircra!t to roll 'ac4 leel + sideslip sta'ilit#. &ncreasing dihedral increases sideslip sta'ilit# and #ou will then hae to hold on more aileron to hold #our attitude. Longitudinal sta'ilit# is directl# in!luenced '# centre o! grait# position. &! an airplane has poor longitudinal sta'ilit# in !light what can 'e done to increase the sta'ilit# &ncrease sta'ili8er sur!ace area. he a!t "; limit can 'e determined '# the minimum accepta'le static longitudinal sta'ilit#. , !orward "; shi!t decreases longitudinal manoeura'ilit#. he dihedral construction o! an aircra!t wing proides Lateral sta'ilit# a'out the longitudinal axis. Positie static longitudinal sta'ilit# means that a nose+down moment occurs a!ter encountering an up+gust. ,!ter an aeroplane has 'een trimmed the stic4 position sta'ilit# will 'e unchanged. -tic4 !orce per g is dependent on cg location. ,n aeroplane:s sideslip angle is de!ined as the angle 'etween the speed ector and the plane o! s#mmetr#.
•
•
•
• •
•
•
•
•
•
•
•
•
•
•
•
•
•
•
•
•
he e!!ect o! the wing downwash on the static longitudinal sta'ilit# o! an aeroplane is negatie. he stic4 !orce per g must hae 'oth an upper and lower limit in order to ensure accepta'le control characteristics. Cpward de!lection o! a trim ta' in the longitudinal control results in the stic4 position sta'ilit# remaining constant. , 'o' weight and a down spring hae the same e!!ect on the stic4 !orce sta'ilit#. 1hat is the e!!ect o! eleator trim ta' adustment on the static longitudinal sta'ilit# o! an aeroplane Go e!!ect. 1hich part o! an aeroplane proides the greatest positie contri'ution to static longitudinal sta'ilit# he hori8ontal tailplane. -tatic lateral sta'ilit# should not 'e too large 'ecause too much aileron de!lection would 'e re*uired in a crosswind landing. 1hen the "; is close to the !orward limit$ Ver# high stic4 !orces are re*uired in pitch 'ecause the aircra!t is er# sta'le. he e!!ect o! ach trim on stic4 !orces !or power operated controls$ &s to maintain the re*uired stic4 !orce gradient. &! the aircra!t is properl# loaded the "; the neutral point and the manoeure point will 'e in the order gien !orward to a!t$ "; neutral point manoeure point. , negatie contri'ution to the static longitudinal sta'ilit# o! conentional et transport aeroplanes is proided '#$ the !uselage. ,n# o! the design !eatures that increases the static lateral sta'ilit# '# increasing the li!t on the low wing will add to the static directional sta'ilit#. , straight wing with no dihedral will not hae much o! a change in li!t on the low wing during a sideslip so it will not hae the increase in drag to #aw it into the relatie air!low. "ontrol sur!ace !lutter can 'e eliminated '#$ mass 'alancing o ! the control sur!ace. he positie manoeuring limit load !actor !or a light ae roplane in the utilit# categor# in the clean con!iguration is$ 7.7. he relationship 'etween the stall speed V- and V, (D,-) !or a large transport aeroplane can 'e expressed in the !ollowing !ormula$ V, T= V- 9 -X(2.B). For most et transport aeroplanes the maximum operating limit speed VO$ is replaced '# O at higher altitudes. he stall speed line in the manoeuring load diagram runs through a point where the$ speed = V- load !actor = N0. 1hat can happen to the aeroplane structure !l#ing at a speed ust exceeding V, &t ma# su!!er permanent de!ormation i! the eleator is !ull# de!lected upwards. ,ircra!t designers could 'uild !or a higher sa!e g limit i! the# wanted to 'ut cost and weight considerations usuall# mean that 2.B g is 'oth the minimum and the maximum. Load !actor is increased '#$ upward gusts.
•
•
•
he extreme right limitation !or 'oth gust and manoeure diagrams is created '# the speed$ V3. he gust load !actor due to a ertical upgust increases when$ the gradient o! the "L+alpha graph increases. ;ust Load Factors ar# depending on altitude mass/weight speed and the slope o! the "L+alpha cure. 1ith increasing altitude and increasing mass/weight the gust load !actor will decrease. 1ing loading is the ratio o! all+up weight/wing area so li4e mass/weight this will also result in a reduced gust load !actor. "onersel# i! speed is increased then the e!!ect o! an upgust will 'e more seere and lead to an increased gust load !actor.
•
•
•
•
• •
• • •
•
• •
•
•
•
• •
•
1hich !actor should 'e ta4en into account when determining V, he limit load !actor. ,ileron !lutter can 'e caused '#$ c#clic de!ormations generated '# aerod#namic inertial and elastic loads on the wing. 0. &ncreasing the aspect+ratio o! the wing will increase the gust load !actor.2. &ncreasing the speed will increase the gust load !actor. &. ,ero+elastic coupling a!!ects !lutter characteristics. &&. he ris4 o! !lutter increases as &,- increases. ,ileron reersal can 'e caused '#$ wisting o! the wing a'oe reersal speed. ,ll gust lines in the gust load diagram originate !rom a point where the speed = load !actor = N0 &! clim'ing at VO it is possi'le to exceed O. VO$ should 'e not greater than V". 1hich o! these statements concerning !light in tur'ulence is correct V, is the recommended tur'ulence penetration air speed. he manoeuring speed V, expressed as indicated airspeed o! a transport aeroplane$ depends on aeroplane mass and pressure altitude. he gust limit load !actor can 'e higher than the manoeuring limit load !actor. ,ssuming &-, conditions which statement with respect to the clim' is correct ,t constant &,- the ach num'er increases. he positie manoeuring limit load !actor !or a large transport aeroplane with !laps extended is$ 2.. he stall speed lines in the manoeuring load diagram originate !rom a point where the$ speed = load !actor = . @ow can wing !lutter 'e preented ?# locating mass in !ront o! the torsion axis o! the wing. V, is$ the maximum speed at which maximum eleator de!lection up is allowed. he load !actor in tur'ulence ma# !luctuate a'oe and 'elow 0 and can een 'ecome negatie. ass+'alancing o! control sur!aces is used to$ preent !lutter o! control sur!aces.
•
•
•
•
•
• •
•
• •
• •
•
•
•
•
•
•
•
•
For an aeroplane with one !ixed alue o! V, the !ollowing applies. V, is$ the speed at which the aeroplane stalls at the manoeuring limit load !actor at O1. he signi!icance o! V, !or et transport aeroplanes is reduced at high cruising altitudes 'ecause$ 'u!!et onset limitations normall# 'ecome limiting. Gow on entering tur'ulence a gust will change the angle o! attac4< the change in "L !or the swept wing will 'e less than !or the straight wing. -wept wings are less sensitie to gusts. "ontrol sur!ace !lutter$ &s a destructie i'ration that must 'e damped out within the !light enelope. Flight in seere tur'ulence ma# lead to a stall and/or structural limitations 'eing exceeded. 1hat is the primar# input !or an arti!icial !eel s#stem &,-. he eleator de!lection re*uired !or a gien manoeure will 'e$ larger at low &,when compared to high &,- larger !or a !orward "; position when compared to an a!t position. Yaw is !ollowed '# roll 'ecause the$ #awing motion generated '# rudder de!lection causes a speed increase o! the outer wing which increases the li!t on that wing so that the aeroplane starts to roll in the same direction as the #aw. Out'oard ailerons (i! present) are normall# used$ in low speed !lights onl#. he !orward "; limit is mainl# determined '# the amount o! pitch control aaila'le !rom the eleator. he centre o! grait# moing a!t will$ increase the eleator up e!!ectieness. 1hen the cg position is moed !orward the eleator de!lection !or a manoeure with a gien load !actor greater than 0 will 'e$ larger. -tic4 !orces proided '# an eleator !eel s#stem depend on$ eleator de!lection d#namic pressure. 1hich 4ind o! ::ta':: is commonl# used in case o! manual reersion o! !ull# powered !light controls -ero ta'. 1hich statement is correct a'out a spring ta' ,t high &,- it 'ehaes li4e a sero ta'. 1hat is the e!!ect o! an a!t shi!t o! the centre o! grait# on (0) static longitudinal sta'ilit# and (2) the re*uired control de!lection !or a gien pitch change (0) educes (2) reduces. 1hich three aerod#namic means decrease manoeuring stic4 !orces -ero ta' + horn 'alance + spring ta'. ,n example o! di!!erential aileron de!lection during initiation o! le!t turn is$ Le!t aileron$ BA up. ight aileron$ 2A down. ,n aeroplane is proided with spoilers and ' oth in'oard and out'oard ailerons. oll control during cruise is proided '#$ in'oard ailerons an d roll spoilers. &n straight and leel !light as speed is reduced$ the eleator is de!lected !urther upwards and the trim ta' !urther downwards.
•
•
•
•
•
•
• •
•
•
•
•
•
•
•
•
•
•
•
1hen roll spoilers are extended the part o! the wing on which the# are mounted$ experiences a reduction in li!t which generates the desired rolling moment. &n addition there is a local increase in drag which suppresses aderse #aw. 1hen power assisted controls are used !or pitch control$ a part o! the aerod#namic !orces is still !elt on the column. he pitch angle is de!ined as the angle 'etween the$ longitudinal axis and the hori8ontal plane. ,ileron de!lection causes a rotation around the longitudinal axis '#$ changing the wing cam'er and the two wings there!ore produce di!!erent li!t alues resulting in a moment a'out the longitudinal axis. ,eroplane manoeura'ilit# decreases !or a gien control sur!ace de!lection when$ &,- decreases. For a gien eleator de!lection aeroplane longitudinal manoeura'ilit# increases when$ the "; moes a!t. 3i!!erential aileron de!lection$ e*uals the drag o! the right and le!t aileron. Dxamples o! aerod#namic 'alancing o! control sur!aces are$ sero ta' spring ta' seal 'etween the wing trailing edge and the leading edge o! control sur!ace. ,n adantage o! locating the engines at the rear o! the !uselage in comparison to a location 'eneath the wing is $ less in!luence o! thrust changes on longitudinal control. ,n aeroplane has a sero ta' controlled eleator. 1hat will happen i! the eleator ams during !light Pitch control sense is reersed. , horn 'alance in a control s#stem has the !ollowing purpose$ to decrease stic4 !orces. 1hat is the position o! the eleator in relation to the trimma'le hori8ontal sta'iliser o! a power assisted aeroplane that is in trim he position depends on speed the position o! slats and !laps and the position o! the centre o! grait#. 1hen a et transport aeroplane ta4es o!! with the "; at the !orward limit and the trimma'le hori8ontal sta'iliser (@-) is positioned at the maximum allowa'le nose down position !or ta4e+o!!$ rotation will re*uire a higher than normal stic4 !orce. 1hen !lutter damping o! control sur!aces is o'tained '# mass 'alancing these weights will 'e located with respect to the hinge o! the control sur!ace$ in !ront o! the hinge. &n a di!!erential aileron control s#stem the control sur!aces hae a larger upward than downward maximum de!lection. @ow does positie cam'er o! an aero!oil a!!ect static longitudinal sta'ilit# &t has no e!!ect 'ecause cam'er o! the aero!oil produces a constant pitch down moment coe!!icient independent o! angle o! attac4. ,n aeroplane:s 'an4 angle is de!ined as the angle 'etween its$ lateral axis and the hori8ontal plane. Low speed pitch+up can 'e caused '# a signi!icant thrust$ increase with podded engines located 'eneath a low+mounted wing. ,rti!icial !eel is re*uired$ with !ull# powered !light controls.
•
•
•
•
• •
•
•
•
•
•
•
• •
•
•
•
•
•
One adantage o! mounting the hori8ontal tailplane on top o! the ertical !in is$ to improe the aerod#namic e!!icienc# o! the ertical !in. , primar# stop is mounted on an eleator control s#stem in order to$ estrict the range o! moement o! the eleator. 1hen ice is present on the sta'ili8er de!lection o! !laps ma# cause$ he sta'ili8er to stall and a ertical die. he !ollowing is true concerning a 'alance ta'. &t is$ a !orm o! aerod#namic 'alance. he inputs to the X !eel unit are !rom$ Pitot and static pressures. he reasons !or haing a trim s#stem on powered assisted !l#ing controls is$ Dna'les the stic4 !orce to 'e reduced to 8ero. 3e!lecting the eleator up when the trim ta' is in neutral will cause the ta' to$ emain in line with the eleator. -ome airplanes hae spring ta's mounted into the control s#stem. his is to proide$ , reduction in the pilots6 e!!ort to moe the controls against high air loads. he :slipstream e!!ect: o! a propeller is most prominent at$ low airspeeds with high power setting. Fixed+pitch propellers are usuall# designed !or maximum e!!icienc# at$ cruising speed &! the propeller pitch o! a windmilling propeller is increased during a glide at constant &,- the propeller drag in the direction o! !light will$ decrease and the rate o! descent will decrease. 3uring a glide with idle power and constant &,- i! the P leer o! a constant speed propeller is pulled 'ac4 !rom its normal cruise position the propeller pitch will$ increase and the rate o! descent will decrease. , windmilling propeller$ produces drag instead o! thrust. he di!!erence 'etween a propeller:s 'lade angle and its angle o! attac4 is called$ the helix angle. he re!erence section o! a propeller 'lade with radius is usuall# ta4en at a distance !rom the propeller axis e*ual to$ .B . For a !ixed+pitch propeller designed !or cruise the angle o! attac4 o! each 'lade measured at the re!erence section$ is optimum when the aircra!t is in a sta'ili8ed cruising !light. "onstant+speed propellers proide a 'etter per!ormance than !ixed+pitch propellers 'ecause the#$ produce an almost maximum e!!icienc# oer a wider speed range. Propeller e!!icienc# ma# 'e de!ined as the ratio 'etween$ usa'le (power aaila'le) power o! the propeller and sha!t power.
D!!icienc# is power out compared to power in so #ou need an aero!oil design that has a good li!t/drag ratio. , high aspect 'lade (long and narrow) will gie #ou low
induced drag and so it needs low engine power to oercome drag. ,s propellers operate at 'ig angles o! attac4 to produce thrust a low induced drag will gie #ou a good e!!icienc#. &! #ou loo4 at #our light aircra!t that is exactl# what the# hae. Fairl# long narrow 'lades (high aspect ratio) and good e!!icienc# so #ou don6t need a 'ig engine to drie them. Cn!ortunatel# the# don:t moe a lot o! air 'ac4wards so although the# are !ine i! the aircra!t weight is low #ou need more disc solidit# i! #ou need more thrust. he minute #ou increase disc solidit# (with wider 'lades) #ou are reducing aspect ratio and losing e!!icienc#. You will need a stronger engine to oercome the higher drag. Power a'sorption is the same as disc solidit# in that it reduces e!!icienc#. &! #ou hae a power!ul engine #ou are wasting it i! #ou put a high aspect ratio (low disc solidit#) propeller on it 'ecause #ou could drie it at idle. You might as well use the power aaila'le '# increasing the disc solidit# and produce more thrust.
•
•
&ncreasing disc solidit# or a'sor'ing more aaila'le power will alwa#s 'e contrar# to e!!icienc#. he designer has to decide what he wants #ou cannot hae 'oth. he angle o! attac4 o! a !ixed pitch propeller 'lade increases when$ P increases and !orward elocit# decreases &ncreasing speed reduces the angle o! attac4 on the prop 'lades. &! #ou had a !ixed pitch prop this would mean that #ou lost thrust and the !aster #ou went the less thrust #ou would hae. ?ut #ou hae a constant speed prop and as the angle o! attac4 reduces the tor*ue drag reduces and the prop P 'egins to go up. he "-C senses this and increases the 'lade angle to get the P sta'ili8ed again. -o #ou settle down at the higher speed with increased 'lade angle.
•
•
•
•
•
•
hin4 @urricane 4 & !ixed pitch wooden prop with a horrendousl# coarse pitch !or high speed. For an# propeller$ thrust is the component o! the total aerod#namic !orce on the propeller parallel to the rotational axis. &! - is the !rontal area o! the propeller disc propeller solidit# is the ratio o!$ the total !rontal area o! all the 'lades to -. &ncreasing the cam'er on propeller 'lades will i! all else is the same$ &ncrease the power a'sorption capa'ilit#. he num'er o! 'lades in a propeller would 'e increased$ o increase power a'sorption capa'ilit#. "ounter rotating propellers hae the e!!ect o!$ "anceling out the tor*ue and g#roscopic e!!ects. he !irst action in the eent o! propeller runawa # (oerspeed conditions) should 'e to$ "lose the throttle.
•
•
•
•
•
•
• •
•
•
•
•
•
•
• •
•
•
•
•
•
•
@ow will the area ratio o! a propeller 'e calculated ,rea o! all propeller 'lades to the total circular sur!ace. 3uring a straight stead# clim' and with the thrust !orce parallel to the !light path$ li!t is the same as during a descent at the same angle and mass. For shallow !light path angles in straight and stead# !light the !ollowing !ormula can 'e used$ sin gamma = /1 + "3/"L. urning motion in a stead# leel co+ordinated turn is created '#$ the centripetal !orce. &n a co+ordinated hori8ontal turn the magnitude o! the centripetal !orce at 7B degrees o! 'an4$ is e*ual to the weight o! the aeroplane. 1hat !actors determine the distance traeled oer the ground o! an aeroplane in a glide he wind and the li!t/drag ratio which changes with angle o! attac4. he speed V"L can 'e limited '# the aaila'le maximum roll rate. For a gien aeroplane which two main aria'les determine the alue o! V"; ,irport eleation and temperature. he airload on the hori8ontal tailplane (tailload) o! an aeroplane in straight and leel cruise !light$ is in general directed downwards and w ill 'ecome less negatie when the c.g. moes a!t. he descent angle o! a gien aeroplane in a stead# wings leel glide has a !ixed alue !or a certain com'ination o!$ (ignore compressi'ilit# e!!ects and assume 8ero thrust) con!iguration and angle o! attac4. &n a stead# straight clim' at clim' angle :gamma: the li!t o! an aeroplane with weight 1 is approximatel#$ 1 9 cos (gamma) 1hen an aeroplane per!orms a straight stead# clim' with a 2H clim' gradient the load !actor is e*ual to$ .UI. 1hich o! the !ollowing parameters can 'e read !rom the para'olic polar diagram o! an aeroplane he minimum glide angle and the parasite drag coe!!icient. &n a straight stead# clim' the thrust must 'e$ greater than the drag 'ecause it must also 'alance a component o! weight. he 'an4 angle in a rate+one turn depends on$ ,-. 1h# is V"; determined with the nosewheel steering disconnected ?ecause the alue o! V"; must also 'e applica'le on wet and/or slipper# runwa#s. he li!t to drag ratio determines the$ hori8ontal glide distance !rom a gien altitude at 8ero wind and 8ero thrust. @ow does V"; change with increasing !ield eleation and temperature 3ecreases 'ecause the engine thrust decreases. ,n aeroplane:s !light path angle is de!ined as the angle 'etween its$ speed ector and the hori8ontal plane. &n a stead# hori8ontal co+ordinated turn$ thrust e*uals drag 'ecause there is e*uili'rium o! !orces along the direction o! !light. ,n aircra!t in !light is a!!ected '# loads. hese ma# 'e classi!ied as$ "ompressie tensile shear and torsional. &n order to clim' with the speed !or maximum clim' rate the aircra!t should 'e !lown with the &,- at which$ he power excess is maximal.
