Aircraft Performance
Module 2
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
Aerodynamics, air data measurements
2
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
Aerodynamics, air data measurements
3
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 ρ A ds F = ma can be transformed into -dp A = ρ A ds dV/dt Can be rewritten as -dp A = ρ A ds/dt dV or dp = - ρ V dV Assuming ρ is constant, integration of the equation gives p + ρ V2/2 = constant
•
(valid for incompressible flow only)
Aerodynamics, air data measurements
4
Incompressible Bernoulli equation (Cont’d) P is the static pressure ρ V2/2 is the dynamic pressure (q) Sum of static and dynamic pressures is the total pressure 2 • ps + ρ V /2 = pt
• ps + q = pt • p 1 + q1 = p2 + q 2
Direct application of the Bernoulli equation is the pitot-static tube
Aerodynamics, air data measurements
5
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 + ρ Vo2/2
Applying Bernoulli equation between points 0 and 2 •
p 2 = po
∆ p = p1 - p2 = ρ Vo2/2
Aerodynamics, air data measurements
6
Speed of sound and Mach number a = speed of sound 0.5 0.5 a = (γ p/ρ ) = (γ RT)
(T= absolute temperature)
∀ γ = ratio of specific heats (constant pressure and constant volume) ∀ γ = Cp/Cv = 1.4 for air ∀ γ and R remain constant in the atmosphere
0.5 a = constant x T
• •
a = a oθ
0.5
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
th • note : 1 knot = 1 nautical mile (nm) / h, 1 nm corresponds to an arc of one minute (1/60 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
Aerodynamics, air data measurements
7
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 = - ρ V dV (presented earlier) • p/ρ γ = constant = C (derived from Boyle’s Law for adiabatic flow) • The two equations above can be combined to give
C 1/ γ p-1/ γ dp = - V dV • Integration of the equation gives
C 1/ γ p-1/ γ -1 /(-1/γ -1) + V2/2 = constant • Can be rewritten as
(γ /(γ -1))p(C/p)1/ γ + V2/2 = constant
Aerodynamics, air data measurements
8
Compressible Bernoulli equation (Cont’d) • Substituting (C/p)1/ γ = 1/ρ , the Bernoulli equation for compressible fluids becomes (γ /(γ -1))p/ρ + V2/2 = constant • The flow equation may be written for any two points in the fluid (γ /(γ -1))p1/ρ 1 + V12/2 = (γ /(γ -1))p2/ρ 2 + V22/2 • Since γ = 1.4 for air, equations can be rewritten as 3.5 p/ρ + V2/2 = constant 3.5 p1/ρ 1 + V12/2 = 3.5 p2/ρ
2
+ V22/2
Aerodynamics, air data measurements
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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 Sonic Supersonic Transonic Hypersonic Hypervelocity
M<1 M=1 M>1 0.80 < M < 1.3 (Approximately) 5 < M < 10 M > 10
Relationships between total and static temperature, density and pressure can be derived from Bernoulli compressible equation for compressible isentropic flow
Aerodynamics, air data measurements
10
Flow relations near the speed of sound (Cont’d) (γ /(γ -1))p1/ρ
1
+ V12/2 = (γ /(γ -1))p2/ρ
+ V22/2
2
Point 1 = reservoir (subscript T for total) : VT = 0
Point 2 = some point in the channel (no subscript) (γ /(γ -1))pT/ρ T = (γ /(γ -1))p/ρ + V2/2
Knowing that a2 = γ p/ρ and aT2 = γ pT/ρ
:
T
aT2 / (γ -1) = a2 /(γ -1) + V2/2
Dividing each side of the equation by a2 (1/(γ -1)) aT2 / a2 = 1/(γ -1) + ½ V2/ a 2
Knowing that aT2 / a2 = TT / T , V/a = M and rearranging : TT / T = 1 + ((γ -1)/2) M 2 = 1 + 0.2 M 2
From pT/p = (ρ ρ
T
T
/ ρ ) (TT / T ) and pT/p = (ρ
T
/ ρ = (1 + ((γ -1)/2) M 2) 1/( γ
pT/p = (1 + ((γ -1)/2) M 2) γ
/( γ -1)
/ρ )γ :
-1)
= (1 + 0.2 M 2) 2.5
= (1 + 0.2 M 2) 3.5
Aerodynamics, air data measurements
11
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 as AR = b/c for a wing with constant chord (rectangular wing) AR = b2/S • Taper ratio (λ ) is the ratio of the tip chord (ct) to the root chord (cr)
λ
= ct / cr
Aerodynamics, air data measurements
12
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
MAC =
Aerodynamics, air data measurements
2 c ∫ db
S
13
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 (Λ ) is the angle between a line perpendicular to the plane of symmetry of the airplane and the quarter chord of each airfoil section
Aerodynamics, air data measurements
14
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 (α ) • 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
L D
α
V
Aerodynamics, air data measurements
15
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 = 0.5ρ V2
(q in lb/ft2 , ρ in slugs/ft3 and V in ft/sec )
q = 1481.3 δ M2
(q in lb/ft2)
or
Under level flight conditions, L = W and CL can then expressed as CL = W / (qS)
Aerodynamics, air data measurements
16
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)
Aerodynamics, air data measurements
17
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 = ρ Vl/µ • • •
l is length from leading edge (ft) µ is the dynamic viscosity (lb-sec/ft2) µ = 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
Aerodynamics, air data measurements
18
Viscosity effects (cont’d)
RN at the transition is approximately 530,000
Aerodynamics, air data measurements
19
Lift and drag Airplane lift and drag vary as a function of α
•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
Aerodynamics, air data measurements
20
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
Aerodynamics, air data measurements
21
Lift and drag (Cont’d) ΔCDcomp (CD – CDREF )
CD = CDP + KCL2 + ΔCDcomp
CL Fixed CL
CDREF CD
M
Aerodynamics, air data measurements
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Lift and drag (Cont’d)
Aerodynamics, air data measurements
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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 δ • -
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 • -
Aerodynamics, air data measurements
24
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
Aerodynamics, air data measurements
25
High-lift Devices (Cont’d)
Effect of slat on air flow Effect of flaps and slats
Aerodynamics, air data measurements
26
High-lift Devices (Cont’d) ΔCDP
Lift and drag increments from various types of trailing edge flaps relative to clean wing
Aerodynamics, air data measurements ΔCL
27
Air data measurements Introduction Airspeed Mach number Altitude Temperature Relationship between flight parameters Angle of attack Typical Pitot-static system Position errors
Aerodynamics, air data measurements
28
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
Aerodynamics, air data measurements
29
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
Aerodynamics, air data measurements
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Airspeed – Calibrated airspeed
Static ports
Total pressure port
Typical Pitot-static probe
Aerodynamics, air data measurements
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Airspeed – Calibrated airspeed (Cont’d)
The compressible Bernoulli equation is the basis for calibrating the airspeed indicator
pT - p = q c = p [ (1 + ((γ -1)/2) (V/a) 2) γ
/( γ -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 + ((γ -1)/2) (Vc/ao) 2) γ q c = po [ (1 + 0.2 (Vc/ao) ) 2
Vc = ao { 5 [ (q c/po + 1)
0.2857
3.5
/( γ -1)
- 1]
- 1] – 1]} 0.5
Aerodynamics, air data measurements
32
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
Aerodynamics, air data measurements
33
Airspeed – Calibrated airspeed (Cont’d)
(qc-q)/qc (%)
(qc-q)/qc
versus Vc
16 14 12 10 8 6 4 2 0
40,000 ft 30,000 ft 20,000 ft 10,000 ft SL 0
50
100
150
200
250
300
350
Vc
Aerodynamics, air data measurements
34
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 ρ
Ve = { [2γ /(γ -1)] (p/ρ o) [(qc/p + 1) ( γ •
o
-1)/ γ
–1]}0.5
Ve from above equation is in ft/sec with pressures in lb/ft2 and ρ
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 ρ
o
Ve 2 = 0.5 ρ V2
Ve = V σ
0.5
True airspeed V is a function of Ve (I.e. qc and p) and σ
Aerodynamics, air data measurements
35
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
Aerodynamics, air data measurements
36
Airspeed – Equivalent airspeed (Cont’d) Delta Vc versus Vc 20 40,000 ft
Delta Vc
15
30,000 ft 20,000 ft
10
10,000 ft
5
SL
0 0
50
100
150
200
250
300
350
Vc
Aerodynamics, air data measurements
37
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 + ∆ V p + ∆ Vi
Note : ∆ Vi is assumed to be zero in our analyses
Aerodynamics, air data measurements
38
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 + ((γ -1)/2) (V/a) 2) γ /( γ -1) - 1] pT - p = q c = p [ (1 + ((γ -1)/2) M 2) γ M = { (2/(γ -1)) [ ( 1 + q c/p )(γ
-1)/ γ
/( γ -1)
- 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
Aerodynamics, air data measurements
39
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
Aerodynamics, air data measurements
40
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
Aerodynamics, air data measurements
41
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 + ((γ -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)
Aerodynamics, air data measurements
42
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 2 • K can be determined by plotting 1/Tt as a function of M / Tt • Slope of fitted test points is equal to –0.2 K or – K/5 • Example is presented on the next page
Aerodynamics, air data measurements
43
Temperature (Cont’d)
Determination of temperature probe recovery factor K
Aerodynamics, air data measurements
44
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 = Vc = ao { 5 [(qc/po + 1) 0.2857
= f (p) – 1]} 0.5
Ve = { 7 (p/ρ o) [(qc/p + 1) 0.2857 V = a { 5 [ (q c/p + 1) 0.2857
–1]}0.5
– 1]} 0.5
M = { 5 [ (qc/p + 1)0.2857 - 1] } 0.5
= f (qc) = f (qc,p) = f (qc,p,a or T) = f (qc,p)
Be careful with units! Aerodynamics, air data measurements
45
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
Aerodynamics, air data measurements
46
Typical Pitot-Static System
Aerodynamics, air data measurements
47
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
Aerodynamics, air data measurements
48
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
Aerodynamics, air data measurements
49
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
Aerodynamics, air data measurements
50
Position errors (Cont’d)
Aerodynamics, air data measurements
51
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
Aerodynamics, air data measurements
52
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 Aerodynamics, air data measurements
53
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)
Aerodynamics, air data measurements
54
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
∀ δ = p/po
hp = (1 - δ
•
1/5.2559
)/6.87535 x 10-6
qc = qic + Cp qic
Vc = ao { 5 [ (q c/po + 1) 0.2857
– 1]} 0.5
Aerodynamics, air data measurements
55
Position errors (Cont’d) Example : Errors associated with Cp = -0.01 for hpI = 5000 ft VI
hp
∆ hp
Vc ∆ Vp
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
Aerodynamics, air data measurements
56
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
Aerodynamics, air data measurements
57
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 kts Altitude error at SL must not exceed 30 ft at speeds < 100 kts
Aerodynamics, air data measurements
58
Position errors (Cont’d) Airspeed error must not exceed 3 knots per 100 knots Airspeed error must not exceed 5 knots at speeds < 166 knots
Aerodynamics, air data measurements
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