Post on 15-Nov-2014
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Aircraft Performance
Module 2
2Aerodynamics, air data measurements
Where are we?
1 : Introduction to aircraft performance, atmosphere
2 : Aerodynamics, air data measurements
3 : Weights / CG, engine performance, level flight
4 : Turning flight, flight envelope
5 : Climb and descent performance
6 : Cruise and endurance
7 : Payload-range, cost index
8 : Take-off performance
9 : Take-off performance
10 : Enroute and landing performance
11 : Wet and contaminated runways
12 : Impact of performance requirements on aircraft design
3Aerodynamics, air data measurements
Aerodynamics
Incompressible Bernoulli equation
Speed of sound and Mach number
Compressible Bernoulli equation
Flow relations near the speed of sound
Airfoil properties
Viscosity effects
Lift and drag
High-lift devices
4Aerodynamics, air data measurements
Incompressible Bernoulli equation Describes variation of pressure and velocity in a streamtube
• F = ma
• Net force : pA - (p + dp)A = -dp A
• Mass of the element is A ds
• F = ma can be transformed into -dp A = A ds dV/dt
• Can be rewritten as -dp A = A ds/dt dV or dp = - V dV
• Assuming is constant, integration of the equation gives
p + V2/2 = constant
• (valid for incompressible flow only)
5Aerodynamics, air data measurements
Incompressible Bernoulli equation (Cont’d) P is the static pressure
V2/2 is the dynamic pressure (q)
Sum of static and dynamic pressures is the total pressure
• ps + V2/2 = pt
• ps + q = pt
• p1 + q1 = p2 + q2
Direct application of the Bernoulli equation is the pitot-static tube
6Aerodynamics, air data measurements
Incompressible Bernoulli equation (Cont’d)
Tube is aligned with the flow
Freestream static pressure and velocity are p0 and V0
At point 1, V1 = 0 (stagnation point)
At point 2, V2 = V0
Applying Bernoulli equation between points 0 and 1
• p1 = po + Vo2/2
Applying Bernoulli equation between points 0 and 2
• p2 = po
p = p1 - p2 = Vo2/2
7Aerodynamics, air data measurements
Speed of sound and Mach number a = speed of sound a = (p/)0.5 = (RT)0.5 (T= absolute temperature)
= ratio of specific heats (constant pressure and constant volume) = Cp/Cv = 1.4 for air and R remain constant in the atmosphere
a = constant x T0.5
a = ao0.5
• ao is the speed of sound under SL ISA conditions (T= 15oC)
• ao = 661.48 knots = 1116.45 ft/sec = 340.28 m/s = 1225.0 km/h
• note : 1 knot = 1 nautical mile (nm) / h, 1 nm corresponds to an arc of one minute (1/60th of a degree) over the earth surface
Mach number (M) is the ratio of local air velocity to local speed of sound
M = V / a
8Aerodynamics, air data measurements
Compressible Bernoulli equation
Air moving at speeds below 200 knots may be treated as an incompressible fluid
At higher speeds, it is necessary to consider the variation of density as the airflow is compressed an another form of the Bernoulli equation must be used• dp = - V dV (presented earlier)
• p/ = constant = C (derived from Boyle’s Law for adiabatic flow)
• The two equations above can be combined to give
C 1/ p-1/ dp = - V dV
• Integration of the equation gives
C 1/ p-1/-1/(-1/ -1) + V2/2 = constant
• Can be rewritten as
(/( -1))p(C/p)1/ + V2/2 = constant
9Aerodynamics, air data measurements
Compressible Bernoulli equation (Cont’d)
• Substituting (C/p)1/ = 1/, the Bernoulli equation for compressible fluids becomes
(/( -1))p/ + V2/2 = constant
• The flow equation may be written for any two points in the fluid
(/( -1))p1/1 + V12/2 = (/( -1))p2/2 + V2
2/2
• Since = 1.4 for air, equations can be rewritten as
3.5 p/ + V2/2 = constant
3.5 p1/1 + V12/2 = 3.5 p2/2 + V2
2/2
10Aerodynamics, air data measurements
Flow relations near the speed of sound Behaviour of fluid flow near the speed of sound is of
primary importance Classification of high speed flight
• Subsonic M < 1
• Sonic M = 1
• Supersonic M > 1
• Transonic 0.80 < M < 1.3 (Approximately)
• Hypersonic 5 < M < 10
• Hypervelocity M > 10
Relationships between total and static temperature, density and pressure can be derived from Bernoulli compressible equation for compressible isentropic flow
11Aerodynamics, air data measurements
Flow relations near the speed of sound (Cont’d) (/( -1))p1/1 + V1
2/2 = (/( -1))p2/2 + V22/2
Point 1 = reservoir (subscript T for total) : VT = 0 Point 2 = some point in the channel (no subscript)
(/( -1))pT/T = (/( -1))p/ + V2/2
Knowing that a2 = p/ and aT2 = pT/T :
aT2 / ( -1) = a2 /( -1) + V2/2
Dividing each side of the equation by a2
(1/( -1)) aT2 / a2 = 1/( -1) + ½ V2/ a 2
Knowing that aT2 / a2 = TT / T , V/a = M and rearranging :
TT / T = 1 + (( -1)/2) M 2 = 1 + 0.2 M 2
From pT/p = (T / ) (TT / T ) and pT/p = (T / ) :
T / = (1 + (( -1)/2) M 2) 1/( -1) = (1 + 0.2 M 2) 2.5
pT/p = (1 + (( -1)/2) M 2) /( -1) = (1 + 0.2 M 2) 3.5
12Aerodynamics, air data measurements
Airfoil properties Physical properties of the wing
• Wingspan (b) is the tip to tip dimension of the wing
• Chord (c) is the distance from the wing leading edge to the trailing edge
• Wing area (S) is the projection of the outline of the plane of the chord
• Aspect ratio (AR) of a wing is defined asAR = b/c for a wing with constant chord (rectangular
wing)AR = b2/S
• Taper ratio () is the ratio of the tip chord (ct) to the root chord (cr)
= ct / cr
13Aerodynamics, air data measurements
Airfoil properties (Cont’d) Physical properties of the wing (Cont’d)
• Mean aerodynamic chord (MAC) is the chord of a section of an imaginary airfoil on the wing which would have force vectors throughout the flight range identical to those of the actual wing
- Can be determined graphically or by integration
S
dbcMAC
2
14Aerodynamics, air data measurements
Airfoil properties (Cont’d) Physical properties of the wing (Cont’d)
• MAC is used as a reference for locating the relative positions of the wing center of lift and the airplane center of gravity (CG)
• Center of lift is normally located at the quarter chord (c/4) of the MAC
• Sweepback () is the angle between a line perpendicular to the plane of symmetry of the airplane and the quarter chord of each airfoil section
15Aerodynamics, air data measurements
Airfoil properties (Cont’d) Aerodynamic properties of the wing
• Pressure distribution around an airfoil in an airflow is a function of the airfoil shape (camber) and the angle of attack ()
• The angle of attack is the relative angle between the freestream velocity (Vo) and the chord (or the fuselage)
• The integration of the pressure distribution around the airfoil can be resolved in two component forces acting at the center of pressure
Lift (L) is perpendicular to the freestream velocity
Drag (D) is parallel to the freestream velocity
V
L
D
16Aerodynamics, air data measurements
Airfoil properties (Cont’d)
L = CLqS
CL is the lift coefficient, CL = L / (qS) , dimensionless
S is the wing area (ft2)
q is the dynamic pressure (lb/ft2)
q = 0.5V2 (q in lb/ft2 , in slugs/ft3 and V in ft/sec )
or
q = 1481.3 M2 (q in lb/ft2)
Under level flight conditions, L = W and CL can then expressed as
CL = W / (qS)
17Aerodynamics, air data measurements
Airfoil properties (Cont’d)
Lift is normally defined in terms of the load factor normal to the flight path Nz
L = NzW or Nz = L / W
Under level flight conditions, Nz = 1.0 (I.e. 1 g)
A more general equation defining CL can be introduced
CL = NzW / (qS)
D = CDqS
CD is the drag coefficient, dimensionless
CD = D / (qS)
18Aerodynamics, air data measurements
Viscosity effects Viscosity is the result of shear forces acting on the fluid, or the tendency of one layer of fluid to
drag along the layer next to it The boundary layer is a finite thickness of fluid next to a surface which is retarded relative to the
free stream velocity Flow in the boundary layer can be laminar or turbulent depending upon
• the smoothness of flow approaching the body, the shape of the body, the surface roughness, the pressure gradient in the direction of flow and the Reynolds number (RN) of the flow (dimensionless)
RN = Vl/ • l is length from leading edge (ft)
• is the dynamic viscosity (lb-sec/ft2)
• = 0.3125 x 10-7 T1.5 / (T + 120) where T is in oK
Friction drag of laminar flow is smaller than friction drag of turbulent flow
19Aerodynamics, air data measurements
Viscosity effects (cont’d)
RN at the transition is approximately 530,000
20Aerodynamics, air data measurements
Lift and drag Airplane lift and drag vary as a function of
•Linear portion of lift curve is where the aircraft normally operates• For performance analysis, it is more practical to look at the variation of CD in terms of CL
21Aerodynamics, air data measurements
Lift and drag (Cont’d) In the range of CL corresponding to normal low speed operation, CD can be closely approximated
by :
CD = A + B CL 2 = CDP + CDI
CDP = parasite drag
CDi = induced drag = CL 2 / ( AR e)
CDi= K CL 2, K is the induced drag factor e is the Oswald efficiency factor
• CL / CD = L / D = Lift to drag ratio• L / D max = max. L/D or min. drag
22Aerodynamics, air data measurements
Lift and drag (Cont’d) Drag increment due to compressibility - ΔCD comp
• Drag increases as a result of separation induced by shock waves during flight at high speed
• Compressibility effects can also happen at relatively slow aircraft speeds (at high CL)
• Drag rise occurs at the onset of shock wave separation - related to critical Mach number (lowest flight speed at which sonic speed is attained on the wing)
ΔCDcomp
CL
CD M
CD = CDP + KCL2 + ΔCD comp
Fixed CL
CD REF
(CD – CD REF)
23Aerodynamics, air data measurements
Lift and drag (Cont’d) Sweep can be used to alleviate compressibility effects
Drag critical Mach number (MCR) is usually defined as the Mach number that causes a drag coefficient increase of 0.0020 (20 drag counts)
Other factors that can alleviate compressibility effects :• Thin supercritical wing section (aft loaded)
• Area ruling
• Vortex generators
24Aerodynamics, air data measurements
Lift and drag (Cont’d) Other factors affecting airplane drag
• During operation with one engine inoperative, additional drag results from - Windmilling engine – CDWM - Data provided by engine manufacturers – typically a
function of M and
- Airplane control deflections and sideslip required to control an asymmetric thrust condition - CDCNTL - function of yawing moment due to thrust
• Drag increment due to deflection of spoilers – CDSP
- In flight or on the ground (ground lift dumpers)- Function of CL and M
• Drag increment due to landing gear - CDLG
- Function of CL
• Deployment of leading edge and trailing edge high-lift devices- Discussed in the next section
25Aerodynamics, air data measurements
High-lift Devices Two types of devices are commonly used to increase CLMAX, the maximum lift
coefficient, and to reduce the stall speed VS
• Trailing edge flaps
• Leading edge slats
Trailing edge flaps provide an increase in camber• Increase in CL at a given angle of attack – increase is essentially proportional to flap
deflection
• Drag increase
• Reduction of the angle of attack at the stall
Leading edge slats provide smoother air on the upper surface of the wing • Slats take high pressure air from under the wing leading edge through a slot to the
upper surface
• Results in greater CLMAX and greater angle of attack at the stall
• Drag increase
Trailing edge flaps and leading edge slats may be used in combination in order to maximize CLMAX
26Aerodynamics, air data measurements
High-lift Devices (Cont’d)
Effect of flaps and slats
Effect of slat on air flow
27Aerodynamics, air data measurements
High-lift Devices (Cont’d)
Lift and drag increments from various types of trailing edge flaps relative to clean wing
ΔCDP
ΔCL
28Aerodynamics, air data measurements
Air data measurements
Introduction
Airspeed
Mach number
Altitude
Temperature
Relationship between flight parameters
Angle of attack
Typical Pitot-static system
Position errors
29Aerodynamics, air data measurements
Introduction
Air data measurements relate atmospheric parameters to the motion of the aircraft
• Airspeed, Mach number, altitude, temperature and angle of attack are important parameters for performance analysis
The objective of this section is to describe :
• The physical principles normally used for these parameters
• The methods of measurement
• The calibration procedures
• The applicable regulations
30Aerodynamics, air data measurements
Airspeed – True airspeed The airspeed V that has been introduced previously is called the true airspeed
• Sometimes also defined as TAS
The true airspeed is the speed of the aircraft relative to the undisturbed air mass. V is the sum of
• The aircraft ground speed Vg (i.e. speed relative to the earth)
• The wind speed vector
Example : An airplane flies in level flight at a ground speed Vg of 500 knots in a tailwind of 50 knots -> V = 450 knots
V has only limited applications operationally.
Other airspeeds must be defined : Calibrated airspeed Equivalent airspeed
Indicated airspeed
31Aerodynamics, air data measurements
Airspeed – Calibrated airspeed
The airspeed indicator uses the pressures obtained from a Pitot-static probe normally located on the forward fuselage
Pitot-static probes are normally used in pairs (one on each side of the fuselage) and are protected from ice accumulation with an electrical heating system
• Static pressures obtained from pilot and co-pilot sides are cross-coupled so as to minimize effect of sideslip on airspeed and altitude indications
Typical Pitot-static probe
Staticports
Total pressureport
32Aerodynamics, air data measurements
Airspeed – Calibrated airspeed (Cont’d) The compressible Bernoulli equation is the basis for calibrating the airspeed indicator
pT - p = q c = p [ (1 + (( -1)/2) (V/a) 2) /( -1)- 1]
The indicator is only driven by the pressure difference (pT – p) or impact pressure (q c) obtained from a pitot-static installation
• Static pressure (p) and speed of sound (a), which is a function of temperature, are not known
• True air speed V can not be related directly to impact pressure
Solution is to define the calibrated airspeed Vc that is based on standard sea level values for p and a :
q c = po [ (1 + (( -1)/2) (Vc/ao) 2) /( -1)- 1]
q c = po [ (1 + 0.2 (Vc/ao) 2) 3.5- 1]
Vc = ao { 5 [ (q c/po + 1) 0.2857 – 1]} 0.5
33Aerodynamics, air data measurements
Airspeed – Calibrated airspeed (Cont’d)
Vc is equal to V under SL/ISA conditions
Flight at constant calibrated airspeed is equivalent to flight at constant q c
q c is close to q during take-off and landing operations at low altitudes (difference is < 2 % typically)
For a given weight, flight at constant Vc at low altitudes ensures that CL and angle-of-attack are nearly constant even if altitude or air density changes• A simple means to maintain a satisfactory margin to the stall
• It would not be the case during flight at constant true airspeed
34Aerodynamics, air data measurements
Airspeed – Calibrated airspeed (Cont’d)
(qc-q)/qc versus Vc
02468
10121416
0 50 100 150 200 250 300 350
Vc
(qc-
q)/
qc
(%
) 40,000 ft
30,000 ft
20,000 ft
10,000 ft
SL
35Aerodynamics, air data measurements
Airspeed – Equivalent airspeed
The equivalent airspeed, Ve, is equal to Vc corrected for adiabatic compressible flow for the particular altitude
Ve, is based on SL ISA density o
Ve = { [2/( -1)] (p/o) [(qc/p + 1) ( -1)/ –1]}0.5
• Ve from above equation is in ft/sec with pressures in lb/ft2 and o in slugs/ft3
Equivalent airspeed is a function of qc and p
Flight at constant Ve is equivalent to flight at constant q
Ve can be defined as the answer to : How fast do I have to travel in SL ISA air to have the same q that I currently have?
0.5 o Ve 2 = 0.5 V2
Ve = V 0.5
True airspeed V is a function of Ve (I.e. qc and p) and
36Aerodynamics, air data measurements
Airspeed – Equivalent airspeed (Cont’d) Ve is not used operationally to fly the aircraft but it is sometimes used for low speed
performance calculations as it results in simpler calculations
• q = Ve 2 / 295.37 (Ve in knots, q in lb/ft2)
• CL = 295.37 L / (Ve 2 S) (Ve in knots)
Ve = Vc - Vc Vc is the compressibility correction
Vc is always positive because qc is greater than q when compressibility effects are present
Vc is equal to 0 at SL (i.e. Ve = Vc at SL)
Vc is less than 1 knot for take-off and landing operations (i.e. altitude less than 10,000 ft and Vc
less than 200 knots Vc ranges between 10-20 knots for typical cruise conditions
Relationship between Vc , Vc and pressure altitude is presented graphically on the next page
37Aerodynamics, air data measurements
Airspeed – Equivalent airspeed (Cont’d)
Delta Vc versus Vc
0
5
10
15
20
0 50 100 150 200 250 300 350
Vc
Del
ta V
c
40,000 ft
30,000 ft
20,000 ft
10,000 ft
SL
38Aerodynamics, air data measurements
Airspeed – Indicated airspeed
The actual airspeed displayed to the pilot is the indicated airspeed, VI, or IAS
VI, is essentially equal to Vc but contains inherent system errors :
• Instrument error Vi (error in instrument calibration)
• Position error Vp (error due to the fact that p and pT are not equal to free stream values, will be detailed later)
VI = Vc - Vp - Vi
Vc = VI + Vp + Vi
Note : Vi is assumed to be zero in our analyses
39Aerodynamics, air data measurements
Mach number
The compressible Bernoulli equation is also the basis for calibrating the Mach number displayed to the pilot (Mach meter)
pT - p = q c = p [ (1 + (( -1)/2) (V/a) 2) /( -1)- 1]
pT - p = q c = p [ (1 + (( -1)/2) M 2) /( -1)- 1]
M = { (2/( -1)) [ ( 1 + q c/p )( -1 )/ - 1 ] } 0.5
M = f (q c , p)
Many aerodynamic effects are function of M
M is independent of static temperature at a given pressure altitude and calibrated airspeed
40Aerodynamics, air data measurements
Altitude
As discussed previously, the altimeter measures static pressure and converts it into an altitude based on equations for the standard atmosphere
The altimeter is connected to the static pressure source and picks up any existing position error hp (error in instrument calibration, pressure leak, …)
• Position error hp will be detailed later
hp = hpI + hp
41Aerodynamics, air data measurements
Temperature
The free air temperature indicator is very important since the indicated temperature has two specific uses associated with performance
• Determination of true air speed V
• Determination of engine pressure ratio (EPR) or engine fan speed (N1) for the required thrust settings
Free air temperature gages are usually operated by an electrical resistance thermometer probe located on the forward portion of the fuselage
42Aerodynamics, air data measurements
Temperature (Cont’d)
Because of the adiabatic temperature rise due to compressibility, the thermometer probe picks up a temperature reading higher than the static temperature
Tt / T = 1 + (( -1)/2) M2 = 1 + 0.2 M2
The equation relating total and static temperatures must be modified to include a temperature probe recovery factor, K, to account for the fact that the probe may not be able to recover the full temperature rise
Tt =T ( 1 + 0.2 K M2)
43Aerodynamics, air data measurements
Temperature (Cont’d)
The probe recovery factor must be determined by flight testing and its value is normally very close to 1.0
• Aircraft must be flown at constant altitude in a stable air mass with constant temperature
• Aircraft is stabilized at various Mach numbers over the operational speed range
• For each stabilized test point, indicated total temperature (TTI) and Mach number (M) are recorded
• K can be determined by plotting 1/Tt as a function of M2 / Tt
• Slope of fitted test points is equal to –0.2 K or – K/5
• Example is presented on the next page
44Aerodynamics, air data measurements
Temperature (Cont’d)
Determination of temperature probe recovery factor K
45Aerodynamics, air data measurements
Relationship between flight parameters
To summarize, all of the data parameters that can be derived Pitot-static and temperature probes are defined in terms of qc, p and T
hp = = f (p)
Vc = ao { 5 [(qc/po + 1) 0.2857 – 1]} 0.5 = f (qc)
Ve = { 7 (p/o) [(qc/p + 1) 0.2857 –1]}0.5 = f (qc,p)
V = a { 5 [ (q c/p + 1) 0.2857 – 1]} 0.5 = f (qc,p,a or T)
M = { 5 [ (qc/p + 1)0.2857 - 1] } 0.5 = f (qc,p)
Be careful with units!
46Aerodynamics, air data measurements
Angle of attack Aircraft angle of attack (AOA) is normally measured with vane-type AOA
transmitters that can be located on the fuselage, on the wing or on a boom
AOA vanes provide AOA information to stall warning / protection systems, flight controls and flight displays
AOA vanes are calibrated on prototype aircraft in order to determine the relation between local AOA at the vane location and aircraft AOA, normally defined relative to the fuselage longitudinal axis
47Aerodynamics, air data measurements
Typical Pitot-Static System
48Aerodynamics, air data measurements
Position errors Total and static pressures sensed by the Pitot-static
system may not be equal to free stream values for various reasons
Position errors are inaccuracies in static and/or total pressures that result in inaccurate airspeed and altitude indications unless corrections are applied
49Aerodynamics, air data measurements
Position errors (Cont’d) Position errors for total and static pressures are defined in terms of pressure coefficients
Static pressure coefficient Cp
Cp = (plocal – p)/qic
Where plocal = local static pressure
p = free stream static pressure
qic = indicated impact pressure (ptlocal- plocal)
ptlocal = local total pressure
Total pressure coefficient Cpt
Cpt = (ptlocal – pt)/qic
Where pt = free stream total pressure
50Aerodynamics, air data measurements
Position errors (Cont’d)
Total pressure can be measured accurately as long as the lips of the Pitot-static probe are very sharp and the local AOA at the probe is less than approximately 25 degrees
Measurement of free stream static is much more difficult
Static pressure varies significantly along the fuselage
Only a few locations where Cp is equal to zero and these locations change with Mach number and AOA
It is not possible to find one location on the aircraft where plocal = p under all flight conditions
51Aerodynamics, air data measurements
Position errors (Cont’d)
52Aerodynamics, air data measurements
Position errors (Cont’d)
Aircraft manufacturers normally install Pitot-static probes where pressure variation is minimum for most normal flight conditions (e.g. Point 2 on figure presented on last page)• Fuselage mounted flush static ports can also be used in combination with a
Pitot probe but flush static ports are more sensitive to skin waviness effects
Other considerations must also be taken into account in order to select Pitot-static probe location• Interference with stall vanes, temperature probe, ice detectors, doors, …
• Skin waviness effects : airframe to airframe variations may be more important in some areas
Calibrations are done on prototype aircraft to determine the position error (static pressure coefficient Cp) for all flight conditions
53Aerodynamics, air data measurements
Position errors (Cont’d)
• Trailing cone is used to measure reference free stream static pressure• Noseboom is used to measure reference total pressure • Reference pressures are compared with ship pressure to determine Cp
• Calibrated pacer aircraft can also be used to determine position error
54Aerodynamics, air data measurements
Position errors (Cont’d)
Total pressure errors are normally negligible• Errors are normally only a concern at high AOA (stall) if probes are
properly located on the airplane
Static pressure errors are affected by different parameters depending on the speed regime
• Low speed regime (M < 0.3): Cp is typically only a function of AOA or CLI, where CLI = (295.37*NZ*W) / (VKIAS^2*S)
• High speed regime (M > 0.3): Cp is typically a function of M but AOA effects may also be present
Cp can be converted in airspeed, altitude and Mach number errors (Vp
, hp and Mp)
55Aerodynamics, air data measurements
Position errors (Cont’d) Typical problem : Determine Vc and hp knowing Cp, indicated airspeed VI and indicated pressure altitude hpI
• qic is determined from VI :
q ic = po [ (1 + 0.2 (VI/ao) 2) 3.5- 1]
• plocal is determined from hpI :
For hpI< 36,089, plocal = po(1 - 6.87535 x 10-6 hpI)5.2559
• Since Cp qic = plocal – p, p = plocal – Cp qic
= p/po
hp = (1 - 1/5.2559)/6.87535 x 10-6
• qc = qic + Cp qic
Vc = ao { 5 [ (q c/po + 1) 0.2857 – 1]} 0.5
56Aerodynamics, air data measurements
Position errors (Cont’d) Example : Errors associated with Cp = -0.01 for hpI = 5000 ft
VI hp hp VcVp
100 kts 4993 ft - 7 ft 99.5 kts -0.5 kts
200 kts 4977 ft - 23 ft 199.0 kts -1.0 kts
300 kts 4952 ft - 48 ft 298.6 kts -1.4 kts
57Aerodynamics, air data measurements
Position errors (Cont’d)
Static pressure error can either be compensated aerodynamically or electronically in order to minimize errors in indicated airspeed and altitude values
• Design of the pitot-static probe can be modified to compensate (in part) the position error (aerodynamic compensation)
• A Static Source Error Correction (SSEC) can be programmed in the Air Data Computer (ADC) to compensate the position error (electronic compensation)
• Electronic compensation is used to compensate position error on most modern aircraft
• SSEC is typically a function of Mach number, AOA and flap position
58Aerodynamics, air data measurements
Position errors (Cont’d)
Residual altitude and airspeed errors, once compensation is applied, must be presented in the AFM
FAR/JAR 25 defines limits for airspeed (FAR 25.1323) and altitude (FAR 25.1325) errors
Altitude error at SL must not exceed 30 ft per 100 ktsAltitude error at SL must not exceed 30 ft at speeds < 100 kts
59Aerodynamics, air data measurements
Position errors (Cont’d)
Airspeed error must not exceed 3 knots per 100 knotsAirspeed error must not exceed 5 knots at speeds < 166 knots