Mattingly J.D., Heiser W.H., Pratt D.T. Aircraft Engine Design

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386 AIRCRAFT ENGINE DESIGN

enter the afterburner at stations 6 and 16, respectively, in Fig. 9.2a. In addition,

the afterburner fuel flow rate that enters between stations 6A and 6.1 is also spec-

ified. When using the AEDsys software for preliminary design, these values are

displayed on the Engine Station Test Results sheet for the particular operating

point chosen for design (see Table 7.E5). The required thermodynamic state of

the combustion product gases exiting the afterburner at station 7 is also fixed. The

radial and axial location of all components of the afterburner assembly in the en-

gine framework will be determined by matching the layout of the fan, turbine, and

variable area nozzles from their respective preliminary designs.

Although the afterburner fuel flow rate and Engine Station Test Results values

listed for stations 6, 16, and 6A are fixed requirements, those values listed for

station 7 are just "first cut" or "bogey" values, which may have to be adjusted in

order to optimize performance. The Engine Station Test Results does not list any

properties whatever for station 6.1. However, a value of total pressure loss from

station 6A to station 7 is obtainable from the Engine Station Test Results sheet,

and this should be regarded as a design target value. This design quantity can be

represented as

Pt7 et6.1 Pt7

Yr6a_ 7 ~'

.....

]TAB D T(AB

(9.121)

Pt6a Pt6a Pt6.1

which is the product of the total pressure ratios across the afterburner diffuser

(stations 6A to 6.1) and the afterburner (stations 6.1 to 7).

9.3.2 Afterburner

Components

9.3.2.1 Mixer. The hot core gas exiting from the low pressure turbine at

station 6 must be mixed with the cooler, comparatively oxygen-rich bypass air at

station 16, in order to provide a gas mixture with sufficiently high temperature to

vaporize the liquid fuel injected and atomized by the spray bars downstream, and

also with sufficiently high temperature and oxygen mol fraction to sustain combus-

tion in the wake of the vee-gutter flameholders located still further downstream. A

schematic representation of the near-constant-area mixer is presented in Fig. 9.35.

A control volume analysis for conservation of mass, momentum, and energy,

presented in Appendix G, determines the state of the mixed gases at station 6A. This

calculation is performed by the AEDsys cycle analysis software, with the results

given on the Engine Station Test Results sheet. However, this control volume

station 16

U " bypass air ........ -~

, ~ ~oregas

--1

station 6

station 6A

L mix

Fig. 9.35 Schematic of afterburner mixer: .... , micromixed region (see Sec. 9.1.4).

DESIGN: COMBUSTION SYSTEMS 387

analysis is based on the assumption that mixing is complete at exit and does not

address the question of how long the mixer must be to accomplish this mixing.

The length required for complete mixing of the core gas and bypass airstreams,

as denoted by

Lmix

in Fig. 9.35, can be estimated by using the shear/mixing layer

equations [Eqs. (9.52-9.60) from Sec. 9.1.4 (see also Figs. 9.18 and 9.19)]. How-

ever, as Fig. 9.35 suggests, the distance required may be excessively long, as one

stream may "mix out" while the other stream still contains unmixed gas. One way

to reduce the length required to mix out is to reduce the annular height of both

passages at mixer inlet, which reduces the initial scale of segregation of the two

mixant streams. 13 This can be done by constant-area ducting (that is, by keeping

the product of cross-sectional radial height H and mean radius

constant as

r m

is increased) both streams to the greatest permissible outer radius prior to mix-

ing. Some other ways are to l) increase the velocity ratio of the two streams [see

Eqs. (9.52-9.54)], 2) purposely imbalance or "un-match" the static pressures in the

two streams so that the higher-pressure stream is pushed laterally into the lower-

pressure stream, and 3) replace the flat-edge splitter plate illustrated in Fig. 9.35

with a "fluted" plate, which can be visualized as a flat plate trailing edge that is

smoothly crimped laterally, similar to hand-crimping the top and bottom layers of

a pie crust. The fluted splitter plate induces vertical velocity components, spatially

alternating upward and downward from both mixant streams, which reduces the

scale of segregation between the two mixants to the pitch of the corrugations in

the fluted splitter plate. Of course, all of the suggested ways to shorten

Lmix

cause

further increases in total pressure loss.

In any case the rational design of the mixer is far too complicated for treatment

at our level of description, 13 and so the properties at stations 6, 16, and 6A listed

on the Engine Station Test Results sheet will simply be taken as given for prelim-

inary design purposes, and the required length can only be guessed or based on

observation of similar designs.

An alternative approach is to not mix the two streams at all. By keeping the two

streams separate until just upstream of the hot section, temperature modulation

for the afterburner can be achieved by having two separately fueled afterburner

systems, one located in the core stream and the other in the unmixed bypass

fan airstream. Stage 1 would be the inner, core stream afterburner, and after the

first stage is lit off, the separately fueled, outer, bypass fan airstream afterburner

could be ignited for stage 2. This in turn leads to flameholding difficulties in both

streams because although the hot core stream is oxygen poor, compared to the

bypass stream, the oxygen-rich bypass stream (air) is cold in comparison with the

core stream.

9.3.2.2 Diffuser. The flow entering the afterburner at station 6A (Fig. 9.2a)

must be slowed to a Mach number that provides a balance between total pressure

loss and afterburner cross-sectional area. The minimum Mach number entering

the combustion zone of the afterburner at station 6.1 is usually fixed by a require-

ment that the diameter of the afterburner section not exceed that of the engine

components located upstream, specifically the fan casing, for case of installation

and removal. A short diffuser length is desired, of course without producing flow

separation, in order to reduce engine weight and length. In augmented turbofan

engines, the diffuser may be combined with a mixer so that a mixed stream enters

388 AIRCRAFT ENGINE DESIGN

Fig. 9.36

station 6A station 6.1

Geometry of afterburner annular flat-wall + dump diffuser.

the combustion section, but the mixer and diffuser will be treated as separate

components for present purposes.

Figure 9.36 shows the geometry and layout of the afterburner diffuser to be

designed. Note that the dump diffuser is defined by truncating the tailcone of

the low-pressure turbine. Extensive use will be made of the concepts and design

equations and principles developed in Secs. 9.1.5 and 9.2.3.

The flow areas and linear dimensions of the annular geometry diffuser shown

in Fig. 9.36 are

A = zr(r2o - r2i) = zr(ro

- ri)(ro

-t- ri)

2rr(ro - ri) (ro + ri)

2zr H

r m

(9.122)

where H =

(ro - ri)

and

= (ro +

)/2,

and for each of the three stations shown,

Al=yr(r 2-1"2)1

Am = Tr(r 2 - r2m) A2 = Jrr~

H1 = (ro - ri)l nm=

(r2 -

rim) H2 -= r2

L = x2 - xl Lm = Xm - Xl La = x2 - xm

(9.123)

The flow area between stations 1 and m is to be designed to vary linearly with

x according to the flat-wall diffuser geometry of Fig. 9.20 and Eq. (9.61),

A(x) _

[rZ(x) - rZ(x)] _

n(x)rm(x) _ 1 -t- rmm (X

--Xl)

2 tan 0 (9.124)

[ro21 -- rT1]

Hlrml

rml

where 0 will be always be set equal to the "sweet spot" value of Fig. 9.22, 0 =

4.5 deg.

By design, the outer radius will also vary linearly between stations 1 and m,

ro(X) = rol + (r2 -- rol ~ (x - x,) Xl < x < Xm

(9.125)

~LmJ

so that the shape of the diffuser centerbody between xl and

x m

is defined by the

DESIGN: COMBUSTION SYSTEMS 389

variation of inner radius of the flow area with x between i and m, from Eq. (9.124),

ri(x) ~r2(x)Zl (l+rmm( -Xl)2 )

.... tan 4.5deg Xl <x <xm (9.126)

7r rm 1 H1

Note that the axial variation of the inner radius

ri(x)

quasi-linear

in x, not

quadratic, because it varies as the square root of a quadratic function.

If the overall area ratio

AR= A2/A 1

is greater than four, the length of the diffuser

centerbody will be the "sweet spot" value

Lm/H1

= 20.

IfAR

is less than four,

then from Eqs. (9.103) and (9.124)

Lm rmlH1

([A~-I] 1)where [A-~I ]

= -- = rIDmAR

rmm

2 tan 4.5 deg

opt opt

(9.127)

Finally, the diffusion efficiency of the combined diffuser is given by Eq. (9.102),

~IDmAR2(1 - [al/am] 2) + 2(AR[A1/Am] -

(9.128)

riD =

(AR 2 -

where, as before,

denotes the overall area ratio

AR = A2/A1.

009 ° from

Eq. (9.68) should be used for 0ore in Eq. (9.128), the total pressure loss can be

calculated from Eq. (9.67), and the tailpipe length

of the combined diffuser

will be

Ld = H2 - Hm.

9.3.2.3 Fuel injection, atomization, and vaporization.

This subject area

is best summarized by Zukoski 3 and is quoted here:

The goal of the fuel injection system is to produce a specified distribution

of fuel vapor in the gas stream entering the afterburner. In most engines, fuel is

introduced in a staged manner so that heat addition rate can be increased gradually

from zero to the desired value. Because ignition, flame stabilization, and flame

spreading are easiest to achieve when the fuel-air ratio is close to the stoichiometric

value, staging is usually produced by adding fuel to successive annular stream

tubes so that the mixture ratio in each tube is nearly stoichiometric. Each stream

tube has its own set of fuel injectors and control system which can be activated

independently.

The most remarkable fact concerning the fuel system for afterburners is

their simplicity. In many engine systems, fuel is supplied to a circular tube which

lies with its axis perpendicular to the gas stream. Fuel is injected into the gas

through small diameter holes located in the sides of the tubes such that the liquid

jet enters the gas stream in a direction perpendicular to the undisturbed flow

direction. The liquid jet penetrates some distance into the gas stream before its

momentum is dissipated. During this penetration process, the air stream tears the

jet apart and produces droplets with diameters of micron size. Heat transfer from

the hot gas stream then vaporizes the droplets.

Given the wide range of values of mass flows of fuel required, it is remark-

able that reasonably thorough mixing of the fuel with the air can be achieved with

390 AIRCRAFT ENGINE DESIGN

this simple injection system. In some recent engines efforts are being made to use

simple variable area injection ports which may possibly give better preparation

of the fuel-air mixture.

The whole area of fuel penetration, atomization, and vaporization is not

well understood from first principles and one of the time consuming parts of an

afterburner development program is to determine the optimum distribution of

locations for injector tubes, injector ports, and port diameters.

A good source of information concerning fuel spray injection, atomization, and

vaporization is given by Chigier. 3°

9. 3.2.4 Ignition. Ignition of the fuel-air mixture in the afterburner is usually

accomplished using a spark or arc igniter or a pilot burner. Once initiated in the

primary stream tube, combustion continues in the wake of the fame stabilizer, and

the process will spread to the rest of the flame stabilizers.

The spark or arc igniter uses a high-energy electric arc to initiate combustion of

the primary stream tube. The igniter is usually placed in the wake of a sheltered

flame stabilizer having its own fuel supply. A stable flame results, and combustion

is initiated behind other flame stabilizers by the mechanism just mentioned.

The pilot burner consists of a pilot zone where a small portion of the inlet

air (usually 10% or less) is burned to stoichiometric temperatures in an enclosed

protected region. The hot gases generated by the pilot are used as an ignition and

stabilizing source for the main fuel injection system.

Afterburning for some turbofan engines is accomplished by adding fuel first

to the core flow near the interface between the core and fan streams, then to the

fan stream, and finally to the rest of the core flow. Afterburning in the fan stream

produces the largest performance gain because of the low temperature of this

stream. However, the fan stream's low temperature makes fuel vaporization and

afterburning very difficult. By adding fuel first to the core flow near its interface

with the fan stream, the resulting afterburning stream can act as a pilot for the

combustion process in the fan stream.

9.3.2.5 Flame stabilization. Twogeneraltypesofflamestabilizingdevices

that have been used in afterburners are shown in Fig. 9.37. These are bluff body

flameholder vee-gutters and piloted burners, where a small piloting heat source is

used to ignite the main fuel flow.

The vee-gutter flameholders have the advantage of low flow blockage and low

total pressure loss. They are simple and lightweight and have a good development

history. The wake of a vee-gutter flameholder is divided into two regions: a recir-

culation zone or bubble and a shear-driven mixing layer, as shown in Fig. 9.2b. The

recirculation zone is characterized by a strong reverse flow and very low chemical

reaction rates. As combustion is essentially complete in the recirculation zone, the

temperature is nearly equal to the adiabatic flame temperature corresponding to the

fuel-air ratio of the approaching stream. The mixing layer is characterized by very

strong shear, steep temperature gradients, and vigorous chemical reaction in the

micromixed reaction zone behind the flame front, which is propagating upstream

against the mean velocity within the mixing layer, as shown in Fig. 9.38. Following

the termination of the mixing layer by disentrainment of the recirculating gases,

the flame front continues to propagate outward towards the walls.

Airflow

DESIGN: COMBUSTION SYSTEMS

391

(a) (b)

Fig. 9.37 Two types of flameholders: a) vee-gutter; and b) piloted burner.

The design goal is to select parameters that will cause the flame region to stand at

a location between

Xm,

the mixing transition point shown in Fig. 9.19 and defined

by Eqs. (9.55) and (9.54), and x =

L/2,

where L is the length of the recirculation

bubble, also as shown in Fig. 9.2b. At about x =

L/2

the mixing layer begins to

break down as a result of disentrainment of gases from the mixing layer into the

recirculation bubble. This is an interesting balancing act because the flame front

must not be permitted to propagate so freely upstream that it "flashes back" to the

spray bars, nor must it be allowed to "blow off' downstream because the mixing

layer and recirculation zone are not long enough.

To find the values of parameters that will provide the needed range for flame

stabilization, the mixing zone will be modeled as a single micromixed well-stirred

reactor (WSR) of lateral thickness equal to the average shear layer thickness

and of axial length equal to one-half the length of the recirculation bubble. The

mean residence time or stay time in this WSR can be estimated by

m pSm(L/2) L

ts "~ -- (9.129)

r'n p~m Uc 2Uc

where Uc is the convective velocity within the mixing layer, as defined by

Eq. (9.55), and where the mass density p has been assumed to be approximately

constant. Further, neglecting the density difference between the two mixing streams,

entrainment flame front

., ~_ ~. disentrainment

Fig. 9.38 Entrainment of reactant gases into, and disentrainment out of, the shear-

driven mixing layer, which is responsible for flameholding in a vee-gutter flameholder.

392 AIRCRAFT ENGINE DESIGN

then

Uc "" U2/2,

and Eq. (9.129) may be rewritten as

(L/2) L

t~ ~ -- (9.130)

(U2/2) U2

The requirement for the flame front to stabilize is therefore that the residence

time or stay time

of the gases flowing through the mixing layer must be greater

than, or equal to, the WSR residence time at blowout

tBo

as defined and discussed

in Sec. 9.1.3.

Because ts >

tso

in the mixing layer is required, then the design requirement

for flame stabilization can be stated as

ts .~ -- > tBo

required (9.131)

Substituting the estimate L --- 4W (Ref. 3) into Eq. (9.131), the stabilization

requirement becomes

ts ~ -- > tBo

required (9.132)

Because both U2 and W depend on the blockage B, approach velocity U1, and

vee-gutter half-angle 0, the requirement Eq. (9.132) must be rewritten in terms of

the controllable design parameters H, D, 0, and U1. Conservation of mass requires

that

U1H=Uz(H-W),

so thatUz=U1 ~ - (1-W/H) (9.133)

which enables Eq. (9.132) to be rewritten as

H4W( W)

ts ~ ~l -H 1 - > tBo

required (9.134)

By inspection of Eq. (9.134), it can be seen that for any given values of channel

height H and approach velocity UI, the residence time in the mixing layer is

maximized when

W/H

= 1/2, so that for optimal flameholding

-- > tso

required, when = 0.5 (9.135)

Note, however, that the optimum is fairly flat. For example, if

W/H

= 0.4

is substituted into Eq. (9.134) instead of the optimal

W/H

= 0.5, then the re-

quired ratio

H~ U1

is increased by only 4%, so that variations in practice can be

expected.

The WSR blowout residence time

tso

must be obtained either from experiment

or from chemical-kinetic WSR modeling, for example, from the AEDsys soft-

ware program KINETX. Required inputs to KINETX for calculating

t~o

are the

pressure, temperature, and composition of the approach gas stream and the ratio

of afterburner fuel flow rate to main gas stream flow rate. All of these parameters

except the gas composition are available from the Engine Station Test Results

DESIGN: COMBUSTION SYSTEMS 393

sheet. The composition of the main gas stream can be obtained from the AEDsys

program EQL.

With U1 given from the Engine Station Test Results sheet,

fro

determined from

the AEDsys program KINETX, and the minimum H determined from Eq. (9.134)

for a single vee-gutter, design choices still have to be made for the vee-gutter

lateral dimension D and half-angle 0, and how many rows (rings) of vee-gutters

and associated spray bars will be required.

Equation (9.87) gives

W/H

as a function of the blockage B =

D/H

and half-

angle 0:

B + (1 - ~)~/B sin(0/2) (9.136)

For any value of

W/H, B

and 0 are related by Eq. (9.136), but neither value is

specified. Because a half-angle of 0 = 15 deg has been widely utilized for vee-

gutter flameholders, that value will be used here. For 0 = 15 deg, and over the

useful range 0.3 <

W/H

< 0.5, Eqs. (9.87) and (9.136) can be represented to

good approximation by

B~ 1.034 -0.1 for 0=15deg, 0.3 < H <0"5 (9.137)

With the optimal

W/H

= 0.5 as input, Eq. (9.137) gives the corresponding value

of blockage B = 0.417, and the gutter lateral width D =

= 0.417H. The

corresponding vee-gutter drag coefficient

= 2.4 is obtained from Eq. (9.84).

These values are typical of currently operational afterburners.

However, for

W/H

= 0.4, Eq. (9.137) gives B = 0.314, and Eq. (9.84) gives

the corresponding

= 1.4. From Eq. (9.85) it can be seen that reducing the

design value of

W/H

from 0.5 to 0.4 reduces the "dry" total pressure loss coef-

ficient from 1.0 to 0.444. This reduction of dry total pressure loss by 55%, at a

cost of only 4% reduction in flameholding stability, is significant and should be

considered during design of the afterburner.

9.3.2.6 Flame spread and afterburner length.

Figure 9.39 illustrates

how the channel height H and flame spread angle determine the length of the

afterburner between stations 6.1 and 7 required for the flame front to consume all

of the reactant gases. Regrettably, there does not appear to be a reliable method

to determine the flame spread angle based on first principles, 3 and so a standard

design value of 17 deg total included angle will be assumed.

Assuming that afterburner a) of Fig. 9.39 has a channel height H that is equal

to or slightly greater than the minimum value required by Eq. (9.135), the

single-gutter afterburner b) has an unnecessarily wide stability margin for flame-

holding, and the twice-as-long flame spreading region results in a much longer

afterburner. Why not make the afterburner even shorter by further reducing the

channel height, so that there would be three or four rows (rings) of flameholders?

Because the vee-gutters would then be too small to stabilize the flame!

In summary, two design guidelines have been established for optimizing both

afterburner flameholding and flame spread: 1) H should be as small as possible in

order to minimize the length of the afterburner (flame spread), and 2) U1 should be

small as possible to maximize residence time in the mixing layer (flameholding.)

394 AIRCRAFT ENGINE DESIGN

(a)

••

(b)

Fig. 9.39 Effect of channel height H and flame spread angle on afterburner length.

Both afterburners have the same area and approach velocity Ul, but the channel height

H of b) is twice that of a). Both have the same, optimum flameholding values

W/H =

0.5, blockage B = 0.417, gutter half-angle 0 = 15 deg, and assumed included angle of

17 deg for flame spread (approximately to scale).

Finally, stealth requirements of modern fighter aircraft may require the flame-

holders be located to hide the high temperature of the turbine exit and thus reduce

the infrared signature. Cool bypass air can be ducted and mixed in such a way as

to cool these flameholders.

9.3.2. 7 Afterburner liner. The afterburner liner is used as a cooling liner

and to improve combustion stability. As a cooling liner, it isolates the very high

temperatures from the outer casing, similar to the liner of the main burner. A film

of cooler air is distributed along the length of the cooling liner that reduces the

metal temperature of this liner and subjects the outer casing of the afterburner to

the afterburner pressure and the temperature of the cooling flow.

The liner is also used as a screech or antihowl liner to prevent extreme high

frequency and amplitude pressure fluctuations resulting from combustion insta-

bility or the unsteady release of thermal energy. This is accomplished by use of

multiple holes along the initial length of the liner. Selective acoustic frequencies

can be dampened by the selection of the proper hole size and distribution and the

geometry of the cavity behind the holes.

9.4 Example Engine Component Design: Combustion Systems

9.4.1 AAF Engine Main Burner Design

Strictly speaking, the main burner should be designed for more than a single

operating point. Specifically, 1) the liner dome should be designed for high-altitude

relight, 2) the primary zone should be sized for idle conditions, 3) the liner air

cooling should be designed for the maximum liner gas temperature, and 4) the

combustor length should be designed for the maximum throughput mass flow rate

condition.*

*Sturgess, G. R., private communication, Dayton, OH, July 2002.

DESIGN: COMBUSTION SYSTEMS 395

As pointed out in Sec. 9.1.7, design of the ignition subsystem is regarded as

too complex for inclusion in this textbook, and so design for high-altitude relight

will not be considered. Of the remaining three considerations, it can be seen from

inspection of Table 7.E4, AAF Engine Main Burner Operation, that the maximum

dynamic pressure condition (sea level, flight Mach number M0 = 1.2) establishes

both the maximum gas temperature and maximum throughput condition. In the

interest of clarity and simplicity, therefore, that condition will be chosen as the

single design point for both the main burner and the afterburner. The corresponding

design interface quantities are presented on the Engine Stations Test Results sheet

(see Table 9.El).

As explained in Sec. 9.2.1, the values listed for station 3.2 are to be regarded

as "first-cut" or "bogey" values, and will have to be checked and/or iterated to

ensure adequate jet penetration and mixing, and to accommodate the swirler.

From the component design of the compressor and turbine, the mean radii at

stations 3.1 and 4 have been established as 9 in. and 7 in., respectively. Using the

Table 9.El AAF engine stations test results at design point (maximum

dynamic pressure)

Results at Engine Stations (AEDsys Ver 3.00) Date: 10/1/2002 6:00:00 AM

Filename: C:\Programs\AEDsys\AAF Data\AAF Final Engine.AED

Engine File Name: C:\Programs\AEDsys\AAF Data\AAF Final Engine.ref

Thrust Scale Factor = 1.0000 Percent Thrust = 100

Altitude = 0 ft

Station m dot gamma Pt Tt Press Mach Velocity Area Area* I

(Ibm/s) (psia) (R) (psia) (ft/s) (ft 2 ) (ft 2 ) (lbf)

0 227.10 1.4000 35.637 668.07 14.696 1.2000 1339.68 2.216 2.151 14145.9

1 227.10 1.4000 35.333 668.07 31.808 0.3904 487.28 3.519 2.169 19557.7

2 227.10 1.4000 33.920 668.07 31.170 0.3496 437.67 4.022 2.260 21140.0

13 227.10 1.4000 83.391 893.29 74.176 0.4125 594.28 1.649 1.063 21805.8

core 117.12 1.4000 83.391 893.29 74.176 0.4125 594.28 0.850 0.548 11245.9

bypass 109.98 1.4000 83.391 893.29 74.176 0.4125 594.28 0.798 0.515 10559.9

2.5 117.12 1.4000 83.391 893.29 76.022 0.3660 529.14 0.938 0.548 12198.7

3 117.12 1.4000 585.114 1659.99 550.205 0.2978 589.47 0.218 0.106 19433.6

3.1 104.24 1.4000 585.114 1659.99 550.205 0.2978 589.47 0.194 0.095 17295.9

3.2 104.24 1.4000 579.263 1659.99 576.537 0.0821 163.90 0.677 0.096 56771.8

MB fuel 2.7017

4 106.94 1.3000 555.859 3138.80 303.347 1.0000 2467.65 0.144 0.144 14494.6

4.1 112.80 1.3000 3042.54

4.4 112.80 1.3000 2389.37

4.5 118.65 1.3000 171.378 2334.78 93.526 1.0000 2128.26 0.448 0.448 13867.4

5 118.65 1.3000 78.129 1983.04 64.783 0.5427 1117.06 1.153 0.906 14878.7

6 118.65 1.3000 78.129 1983.04 71.533 0.3702 770.87 1.549 0.906 18796.4

16 109.98 1.4000 83.391 893.29 71.570 0.4725 677.21 0.719 0.515 7407.6

6A 228.63 1.3360 76.859 1513.71 68.248 0.4249 779.89 2.338 1.536 28516.0

AB fuel 8.7033

7 237.33 1.3000 73.016 3600.00 61.806 0.5113 1421.42 3.466 2.613 41335.2

8 237.33 1.3000 73.016 3600.00 39.847 1.0000 2642.73 2.613 2.613 34489.9

9 237.33 1.3000 70.826 3600.00 14.696 1.7079 4036.90 3.711 2.694 37631.8