Flight path analysis U.K.

In some modelling applications the flight path information is provided not as procedural steps but as coordinates in position and time, usually determined by analysis of radar data. This is discussed in Section 2.7.7 of the main text. In this case the equations presented in this Appendix are used ‘in reverse’; the engine thrust parameters are derived from the aircraft motion rather than vice-versa. In general, once the flight path data has been averaged and reduced to segment form, each segment being classified by climb or descent, acceleration or deceleration, and thrust and flap changes, this is relatively straightforward by comparison with synthesis which often involves iterative processes.

B2ENGINE THRUSTU.K.

The propulsive force produced by each engine is one of five quantities that need to be defined at the ends of each flight path segment (the others being height, speed, power setting and bank angle). Net thrust represents the component of engine gross thrust that is available for propulsion. For aerodynamic and acoustical calculations, the net thrust is referred to standard air pressure at mean sea level. This is known as corrected net thrust, F_n/δ.

This will be either the net thrust available when operating at a specified thrust rating, or the net thrust that results when the thrust-setting parameter is set to a particular value. For a turbojet or turbofan engine operating at a specific thrust rating, corrected net thrust is given by the equation

F_n/δ = E + F · V_c + G_A· h + G_B· h² + H · T

(B-1)

where

F_n

is the net thrust per engine, lbf

is the ratio of the ambient air pressure at the aeroplane to the standard air pressure at mean sea level, i.e., to 101,325 kPa (or 1 013,25 mb) [ref. 1]

F_n/δ

is the corrected net thrust per engine, lbf

V_C

is the calibrated airspeed, kt

is the ambient air temperature in which the aeroplane is operating, °C, and

E, F, G_A, G_B, H

are engine thrust constants or coefficients for temperatures below the engine flat rating temperature at the thrust rating in use (on the current segment of the takeoff/climbout or approach flight path), lb.s/ft, lb/ft, lb/ft², lb/°C. Obtainable from the ANP database.

Data are also provided in the ANP database to allow calculation of non-rated thrust as a function of a thrust setting parameter. This is defined by some manufacturers as engine pressure ratio EPR, and by others as low-pressure rotor speed, or fan speed, N₁ . When that parameter is EPR, equation B-1 is replaced by

F_n/δ = E + F · V_C + G_A · h + G_B · h² + H · T + K₁ · EPR + K₂ · EPR²

(B-2)

where K₁ and K₂ are coefficients, from the ANP database that relate corrected net thrust and engine pressure ratio in the vicinity of the engine pressure ratio of interest for the specified aeroplane Mach number.

When engine rotational speed N₁ is the parameter used by the cockpit crew to set thrust, the generalised thrust equation becomes

(B-3)

where

N₁

is the rotational speed of the engine's low-pressure compressor (or fan) and turbine stages, %

= (T + 273)/288,15, the ratio of the absolute total temperature at the engine inlet to the absolute standard air temperature at mean sea level [ref. 1].

is the corrected low pressure rotor speed, %; and

K₃, K₄

are constants derived from installed engine data encompassing the N₁ speeds of interest.

Note that for a particular aeroplane E, F, G_A, G_B and H in equations B-2 and B-3 might have different values from those in equation B-1.

Not every term in the equation will always be significant. For example, for flat-rated engines operating in air temperatures below the break point (typically 30 °C), the temperature term may not be required. For engines not flat rated, ambient temperature must be considered when designating rated thrust. Above the engine flat rating temperature, a different set of engine thrust coefficients (E, F, G_A, G_B and H) _high must be used to determine the thrust level available. Normal practice would then be to compute F_n /δ using both the low temperature and high temperature coefficients and to use the higher thrust level for temperatures below the flat rating temperature and use the lower calculated thrust level for temperature above the flat rating temperature.

Where only low temperature thrust coefficients are available, the following relationship may be used:

(F_n/δ)_high = F · V_C + (E + H · T_B )·(1 – 0,006 · T)/(1 – 0,006 · T_B )

(B-4)

where

(F_n /δ) _high

high-temperature corrected net thrust (lbf),

T_B

breakpoint temperature (in the absence of a definitive value assume a default value of 30 °C).

The ANP database provides values for the constants and coefficients in equations B-1 to B-4.

For propeller driven aeroplanes, corrected net thrust per engine should be read from graphs or calculated using the equation

F_n/δ = (326 · η · P_p/V_T )/δ

(B-5)

where

is the propeller efficiency for a particular propeller installation and is a function of propeller rotational speed and aeroplane flight speed

V_T

is the true airspeed, kt

P_p

is net propulsive power for the given flight condition, e.g. max takeoff or max climb power, hp

Parameters in equation B-5 are provided in the ANP database for maximum takeoff thrust and maximum climb thrust settings.

True airspeed V_T is estimated from the calibrated airspeed V_C using the relationship

(B-6)

where σ is the ratio of the air density at the aeroplane to the mean sea-level value.

Guidance on operation with reduced takeoff thrust U.K.

Often, aircraft takeoff weights are below maximum allowable and/or the available runway field length exceeds the minimum required with the use of maximum takeoff thrust. In these cases, it is common practice to reduce engine thrust below maximum levels in order to prolong engine life and, sometimes, for noise abatement purposes. Engine thrust can only be reduced to levels that maintain a required margin of safety. The calculation procedure used by airline operators to determine the amount of thrust reduction is regulated accordingly: it is complex and takes into account numerous factors including takeoff weight, ambient air temperature, declared runway distances, runway elevation and runway obstacle clearance criteria. Therefore the amount of thrust reduction varies from flight to flight.

As they can have a profound effect upon departure noise contours, modellers should take reasonable account of reduced thrust operations and, to make best possible provision, to seek practical advice from operators.

If such advice is not available it is still advisable to make some allowance by alternative means. It is impractical to mirror the operators' calculations for noise modelling purposes; nor would they be appropriate alongside the conventional simplifications and approximations which are made for the purposes of calculating long term average noise levels. As a practicable alternative the following guidance is provided. It should be emphasised that considerable research is ongoing in this area and thus, this guidance is subject to change.

Analysis of FDR data has shown that the level of thrust reduction is strongly correlated with ratio of the actual takeoff weight to the Regulated Takeoff Weight (RTOW), down to a fixed lower limit(1); i.e.

F_n/δ = (F_n/δ)_max · W/W_RTOW

(B-7)

where (F_n /δ) _max is the maximum rated thrust, W is the actual gross take-off weight and W_RTOW is the Regulated Takeoff Weight.

The RTOW is the maximum takeoff weight that can be safely used, whilst satisfying takeoff field length, engine-out and obstacle requirements. It is a function of the available runway length, airfield elevation, temperature, headwind, and flap angle. This information can be obtained from operators and should be more readily available than data on actual levels of reduced thrust. Alternatively, it may be computed using data contained in aircraft flight manuals.

Reduced Climb Thrust U.K.

When employing reduced take-off thrust, operators often, but not always, reduce climb thrust from below maximum levels(2). This prevents situations occurring where, at the end of the initial climb at take-off thrust, power has to be increased rather than cut back. However, it is more difficult to establish a rationale for a common basis here. Some operators use fixed detents below maximum climb thrust, sometimes referred to as Climb 1 and Climb 2, typically reducing climb thrust by 10 and 20 percent respectively relative to maximum. It is recommended that whenever reduced takeoff thrust is used, climb thrust levels also be reduced by 10 percent.

B3VERTICAL PROFILES OF AIR TEMPERATURE, PRESSURE, DENSITY AND WINDSPEEDU.K.

For the purposes of this document, the variations of temperature, pressure and density with height above mean sea level are taken to be those of the International Standard Atmosphere. The methodologies described below have been validated for aerodrome altitudes up to 4 000 ft above sea level and for air temperatures up to 43 °C (109 °F).

Although, in reality, mean wind velocity varies with both height and time, it is not usually practicable to take account of this for noise contour modelling purposes. Instead, the flight performance equations given below are based on the common assumption that the aeroplane is heading directly into a (default) headwind of 8 kt at all times — regardless of compass bearing (although no explicit account of mean wind velocity is taken in sound propagation calculations). Methods for adjusting the results for other headwind speeds are provided.

B4THE EFFECTS OF TURNSU.K.

The remainder of this appendix explains how to calculate the required properties of the segments joining the profile points s,z that define the two-dimensional flight path in the vertical plane above the ground track. Segments are defined in sequence in the direction of motion. At the end of any one segment (or at the start of roll in the case of the first for a departure) where the operational parameters and the next procedural step are defined, the need is to calculate the climb angle and track distance to the point where the required height and/or speed are reached.

If the track is straight, this will be covered by a single profile segment, the geometry of which can then be determined directly (albeit sometimes with a degree of iteration). But if a turn starts or ends, or changes in radius or direction, before the required end-conditions are reached, a single segment would be insufficient because the aircraft lift and drag change with bank angle. To account for the effects of the turn on the climb, additional profile segments are required to implement the procedural step — as follows.

The construction of the ground track is described in Section 2.7.13 of the text. This is done independently of any aircraft flight profile (although with care not to define turns that could not be flown under normal operating constraints). But as the flight profile — height and speed as a function of track distance — is affected by turns so that the flight profile cannot be determined independently of the ground track.

To maintain speed in a turn the aerodynamic wing lift has to be increased, to balance centrifugal force as well as the aircraft weight. This in turn increases drag and, consequently the propulsive thrust required. The effects of the turn are expressed in the performance equations as functions of bank angle ε which, for an aircraft in level flight turning at constant speed on a circular path, is given by

			(B-8)
where	V	is the groundspeed, kt
where	r	is the turn radius, ft
and	g	is the acceleration due to gravity, ft/s²

All turns are assumed to have a constant radius and second-order effects associated with non-level flight paths are disregarded; bank angles are based on the turn radius r of the ground track only.

To implement a procedural step a provisional profile segment is first calculated using the bank angle ε at the start point — as defined by equation B-8 for the track segment radius r. If the calculated length of the provisional segment is such that it does not cross the start or end of a turn, the provisional segment is confirmed and attention turns to the next step.

But if the provisional segment crosses one or more starts or ends of turns (where ε changes)(3), the flight parameters at the first such point are estimated by interpolation (see Section 2.7.13), saved along with its coordinates as end-point values, and the segment truncated. The second part of the procedural step is then applied from that point — once more assuming provisionally that it can be completed in a single segment with the same end conditions but with the new start point and new bank angle. If this second segment then passes another change of turn radius/direction, a third segment will be required — and so on until the end-conditions are achieved.

Approximate method U.K.

It will be apparent that accounting fully for the effects of turns, as described above, involves considerable computational complexity because the climb profile of any aircraft has to be calculated separately for each ground track that it follows. But changes to the vertical profile caused by turns usually have a markedly smaller influence on the contours than the changes of bank angle, and some users may prefer to avoid the complexity — at the cost of some loss of precision — by disregarding the effects of turns on profiles while still accounting for the bank angle in the calculation of lateral sound emission (see Section 2.7.19). Under this approximation profile points for a particular aircraft operation are calculated once only, assuming a straight ground track (for which ε = 0).

B5TAKEOFF GROUND ROLLU.K.

Take-off thrust accelerates the aeroplane along the runway until lift-off. Calibrated airspeed is then assumed to be constant throughout the initial part of the climbout. Landing gear, if retractable, is assumed to be retracted shortly after lift-off.

For the purpose of this document, the actual takeoff ground-roll is approximated by an equivalent take-off distance (into a default headwind of 8 kt), s_TO8 , defined as shown in Figure B-1, as the distance along the runway from brake release to the point where a straight line extension of the initial landing-gear-retracted climb flight path intersects the runway.

Figure B-1

Equivalent takeoff distance

On a level runway, the equivalent takeoff ground-roll distance s_TO8 in feet is determined from

(B-9)

where

B₈

is a coefficient appropriate to a specific aeroplane/flap-deflection combination for the ISA reference conditions, including the 8-knot headwind, ft/lbf

is the aeroplane gross weight at brake release, lbf

is the number of engines supplying thrust.

Note: U.K.

Since equation B-9 accounts for variation of thrust with airspeed and runway elevation, for a given aeroplane the coefficient B₈ depends only on flap deflection.

For headwind other than the default 8 kt, the takeoff ground-roll distance is corrected by using:

(B-10)

where

S_TOw

is the ground-roll distance corrected for headwind w, ft

V_C

(in this equation) is the calibrated speed at takeoff rotation, kt

is the headwind, kt

The takeoff ground-roll distance is also corrected for runway gradient as follows:

(B-11)

where

S_TOG

is the ground-roll distance (ft) corrected for headwind and runway gradient,

is the average acceleration along the runway, equal to , ft/s²

G_R

is the runway gradient; positive when taking-off uphill

B6CLIMB AT CONSTANT SPEEDU.K.

This type of segment is defined by the aeroplane's calibrated airspeed, flap setting, and the height and bank angle at its end, together with the headwind speed (default 8 kt). As for any segment, the segment start parameters including corrected net thrust are put equal to those at the end of the preceding segment — there are no discontinuities (except of flap angle and bank angle which, in these calculations, are allowed to change in steps). The net thrusts at the segment end are first calculated using the appropriate equation from B-1 to B-5. The average geometric climb angle g (see Figure B-1) is then given by

(B-12)

where the over-bars denote mid-segment values (= average of start-point and end-point values — generally the mid-segment values) and

is a speed-dependent constant equal to 1,01 when V_C ≤ 200 kt or 0,95 otherwise. This constant accounts for the effects on climb gradient of climbing into an 8-knot headwind and the acceleration inherent in climbing at constant calibrated airspeed (true speed increases as air density diminishes with height).

is the ratio of the aeroplane's drag coefficient to its lift coefficient appropriate to the given flap setting. The landing gear is assumed to be retracted.

Bank angle, radians

The climb angle is corrected for headwind w using:

(B-13)

where γ_w is the average climb angle corrected for headwind.

The distance that the aeroplane traverses along the ground track, Δs, while climbing at angle γ _w , from an initial altitude h ₁ to a final altitude h ₂ is given by

(B-14)

As a rule, two distinct phases of a departure profile involve climb at constant airspeed. The first, sometime referred to as the initial climb segment is immediately after lift-off, where safety requirements dictate that the aeroplane is flown at a minimum airspeed of least the takeoff safety speed. This is a regulated speed and should be achieved by 35 ft above the runway during normal operation. However, it is common practice to maintain an initial climb speed slightly beyond the takeoff safety speed, usually by 10-20 kt, as this tends to improve the initial climb gradient achieved. The second is after flap retraction and initial acceleration, referred to as continuing climb.

During the initial climb, the airspeed is dependent on the takeoff flap setting and the aeroplane gross weight. The calibrated initial climb speed V_CTO is calculated using the first order approximation:

(B-15)

where C is a coefficient appropriate to the flap setting (kt/√lbf), read from the ANP database.

For continuing climb after acceleration, the calibrated airspeed is a user input parameter.

B7POWER CUTBACK (TRANSITION SEGMENT)U.K.

Power is reduced, or cut back, from take-off setting at some point after takeoff in order to extend engine life and often to reduce noise in certain areas. Thrust is normally cut back during either a constant speed climb segment (Section B6) or an acceleration segment (Section B8). As it is a relatively brief process, typically of only 3-5 seconds′ duration, is it modelled by adding a ‘transition segment’ to the primary segment. This is usually taken to cover a horizontal ground distance of 1 000 ft (305 m).

Amount of thrust reduction U.K.

In normal operation the engine thrust is reduced to the maximum climb thrust setting. Unlike the take-off thrust, climb thrust can be sustained indefinitely, usually in practice until the aeroplane has reached its initial cruise altitude. The maximum climb thrust level is determined with equation B-1 using the manufacturer supplied maximum thrust coefficients. However, noise abatement requirements may call for additional thrust reduction, sometimes referred to as a deep cutback. For safety purposes the maximum thrust reduction is limited(4) to an amount determined by the performance of the aeroplane and the number of engines.

The minimum ‘reduced-thrust’ level is sometimes referred to as the engine-out ‘reduced thrust’:

(B-16)

where

δ₂

is the pressure ratio at altitude h₂

G′

is the engine-out percentage climb gradient:

= 0 % for aeroplanes with automatic thrust restoration systems; otherwise,
= 1,2 % for 2-engine aeroplane
= 1,5 % for 3-engine aeroplane
= 1,7 % for 4-engine aeroplane

Constant speed climb segment with cutback U.K.

The climb segment gradient is calculated using equation B-12, with thrust calculated using either B-1 with maximum climb coefficients, or B-16 for reduced thrust. The climb segment is then broken into two sub-segments, both having the same climb angle. This is illustrated in Figure B-2.

Figure B-2

Constant speed climb segment with cutback (illustration — not to scale)

The first sub-segment is assigned a 1 000 ft (304 m) ground distance, and the corrected net thrust per engine at the end of 1 000 ft is set equal to the cutback value. (If the original horizontal distance is less than 2 000 ft, one half of the segment is used to cutback thrust.) The final thrust on the second sub-segment is also set equal to the cutback thrust. Thus, the second sub-segment is flown at constant thrust.

B8ACCELERATING CLIMB AND FLAP RETRACTIONU.K.

This usually follows the initial climb. As for all flight segments, the start-point altitude h₁ , true airspeed V_T1 , and thrust (F_n /δ)₁ are those from the end of the preceding segment. The end-point calibrated airspeed V_C2 and the average climb rate ROC are user inputs (bank angle ε is a function of speed and radius of turn). As they are interdependent, the end altitude h₂ , end true airspeed V_T2 , end thrust (F_n /δ)₂ and segment track length Δs have to be calculated by iteration; the end altitude h ₂ is guessed initially and then recalculated repeatedly using equations B-16 and B-17 until the difference between successive estimates is less than a specified tolerance, e.g. one foot. A practical initial estimate is h ₂ = h ₁ + 250 feet.

The segment track length (horizontal distance covered) is estimated as:

(B-17)

where

0,95

is a factor to account for effect of 8 kt headwind when climbing at 160 kt

is a constant to convert knots to ft/sec = 1,688 ft/s per kt

V_T2

= true airspeed at segment end, kt:

where σ₂ = air density ratio at end altitude h ₂

α_max

= maximum acceleration in level flight (ft/s²)

= climb gradient

where ROC = climb rate, ft/min

Using this estimate of Δs, the end altitude h ₂′ is then re-estimated using:

h₂′ = h ₁ + s · G/0,95

(B-18)

As long as the error is outside the specified tolerance, the steps B-17 and B-18 are repeated using the current iteration segment-end values of altitude h ₂, true airspeed V_T2 , corrected net thrust per engine (F_n /δ)₂. When the error is within the tolerance, the iterative cycle is terminated and the acceleration segment is defined by the final segment-end values.

Note: U.K.

If during the iteration process (a_max – G·g) < 0,02 g, the acceleration may be too small to achieve the desired V_C2 in a reasonable distance. In this case, the climb gradient can be limited to G = a_max /g – 0,02, in effect reducing the desired climb rate in order to maintain acceptable acceleration. If G < 0,01 it should be concluded there is not enough thrust to achieve the acceleration and climb rate specified; the calculation should be terminated and the procedure steps revised(5).

The acceleration segment length is corrected for headwind w by using:

(B-19)

Accelerating segment with cutback U.K.

Thrust cutback is inserted into an acceleration segments in the same way as for a constant speed segment; by turning its first part into a transition segment. The cutback thrust level is calculated as for the constant-speed cutback thrust procedure, using equation B-1 only. Note it is not generally possible to accelerate and climb whilst maintaining the minimum engine-out thrust setting. The thrust transition is assigned a 1 000 ft (305 m) ground distance, and the corrected net thrust per engine at the end of 1 000 ft is set equal to the cutback value. The speed at the end of the segment is determined by iteration for a segment length of 1 000 ft. (If the original horizontal distance is less than 2 000 ft, one half of the segment is used for thrust change.) The final thrust on the second sub-segment is also set equal to the cutback thrust. Thus, the second sub-segment is flown at constant thrust.

B9ADDITIONAL CLIMB AND ACCELERATION SEGMENTS AFTER FLAP RETRACTIONU.K.

If additional acceleration segments are included in the climbout flight path, equations B-12 to B-19 should be used again to calculate the ground-track distance, average climb angle, and height gain for each. As before, the final segment height must be estimated by iteration.

B10DESCENT AND DECELERATIONU.K.

Approach flight normally requires the aeroplane to descend and decelerate in preparation for the final approach segment where the aeroplane is configured with approach flap and gear down. The flight mechanics are unchanged from the departure case; the main difference is that the height and speed profile is generally known, and it is the engine thrust levels that must be estimated for each segment. The basic force balance equation is:

(B-20)

Equation B-20 may be used in two distinct ways. First the aeroplane speeds at the start and end of a segment may be defined, along with a descent angle (or level segment distance) and initial and final segment altitudes. In this case the deceleration may be calculated using:

(B-21)

where Δs is the ground distance covered and V ₁ and V ₂ and are the initial and final groundspeeds calculated using

(B-22)

Equations B-20, B-21 and B-22 confirm that whilst decelerating over a specified distance at a constant rate of descent, a stronger headwind will result in more thrust being required to maintain the same deceleration, whilst a tailwind will require less thrust to maintain the same deceleration.

In practice most, if not all decelerations during approach flight are performed at idle thrust. Thus for the second application of equation B-20, thrust is defined at an idle setting and the equation is solved iteratively to determine (1) the deceleration and (2) the height at the end of the deceleration segment — in a similar manner to the departure acceleration segments. In this case, deceleration distance can be very different with head and tail winds and it is sometimes necessary to reduce the descent angle in order to obtain reasonable results.

For most aeroplanes, idle thrust is not zero and, for many, it is also a function of flight speed. Thus, equation B-20 is solved for the deceleration by inputting an idle thrust; the idle thrust is calculated using an equation of the form:

(F_n /δ)_idle = E_idle + F_idle · V_C + G_A,idle · h + G_B,idle · h² + H_idle · T

(B-23)

where (E_idle, F_idle, G_A,idle, G_B,idle and H_idle ) are idle thrust engine coefficients available in the ANP database.

B11LANDING APPROACHU.K.

The landing approach calibrated airspeed, V_CA , is related to the landing gross weight by an equation of the same form as equation B-11, namely

(B-24)

where the coefficient D (kt/√lbf) corresponds to the landing flap setting.

The corrected net thrust per engine during descent along the approach glideslope is calculated by solving equation B-12 for the landing weight W and a drag-to-lift ratio R appropriate for the flap setting with landing gear extended. The flap setting should be that typically used in actual operations. During landing approach, the glideslope descent angle γ may be assumed constant. For jet-powered and multi-engine propeller aeroplanes, γ is typically – 3°. For single-engine, propeller-powered aeroplanes, γ is typically – 5°.

The average corrected net thrust is calculated by inverting equation B-12 using K = 1,03 to account for the deceleration inherent in flying a descending flight path into an 8-knot reference headwind at the constant calibrated airspeed given by equation B-24, i.e.

(B-25)

For headwinds other than 8 kt, average corrected net thrust becomes

(B-26)

The horizontal distance covered is calculated by:

(B-27)

(positive since h₁ > h₂ and γ is negative).

(1)

Airworthiness authorities normally stipulate a lower thrust limit, often 25 percent below maximum.

(2)

To which thrust is reduced after the initial climb at take-off power.

(3)

To avoid contour discontinuities caused by instantaneous changes of bank angle at the junctions between straight and turning flight, sub-segments are introduced into the noise calculations to allow linear transitions of bank angle over the first and last 5° of the turn. These are not necessary in the performance calculations; the bank angle is always given by equation B-8.

(4)

‘Noise Abatement Procedures’, ICAO Document 8168 ‘PANS-OPS’ Vol.1 Part V, Chapter 3, ICAO 2004.

(5)

In either case the computer model should be programmed to inform the user of the inconsistency.

ANNEXU.K.

Appendix B Flight performance calculations

Terms and symbols U.K.

Terms U.K.

Symbols U.K.

B1INTRODUCTIONU.K.

Flight path synthesis U.K.

Flight path analysis U.K.

B2ENGINE THRUSTU.K.

Guidance on operation with reduced takeoff thrust U.K.

Reduced Climb Thrust U.K.

B3VERTICAL PROFILES OF AIR TEMPERATURE, PRESSURE, DENSITY AND WINDSPEEDU.K.

B4THE EFFECTS OF TURNSU.K.

Approximate method U.K.

B5TAKEOFF GROUND ROLLU.K.

Note: U.K.

B6CLIMB AT CONSTANT SPEEDU.K.

B7POWER CUTBACK (TRANSITION SEGMENT)U.K.

Amount of thrust reduction U.K.

Constant speed climb segment with cutback U.K.

B8ACCELERATING CLIMB AND FLAP RETRACTIONU.K.

Note: U.K.

Accelerating segment with cutback U.K.

B9ADDITIONAL CLIMB AND ACCELERATION SEGMENTS AFTER FLAP RETRACTIONU.K.

B10DESCENT AND DECELERATIONU.K.

B11LANDING APPROACHU.K.