The Society shall not be responsible for state69-GT-41
ments or opinions advanced in papers or in discussion at meetings of the Society or of its
Divisions or Sections, or printed in its publications.
Discussion is printed only if the paper is published
in an ASME journal or Proceedings.
$1.50 PER COPY
Released for general publication upon presentation
Copyright © 1969 by ASME
75C TO ASME MEMBERS
The Engine Inlet on the 747
W. S. VIALL
Research Specialist,
The Boeing Co.,
Renton, Wash.
Consideration is given the manner in which inlet throat Mach number, lip losses and
auxiliary passage losses affect the design of a high recovery inlet for high-bypass-ratio fan
engines. Problems relating to auxiliary passage closure are discussed.
Contributed by the Gas Turbine Division for presentation at the Gas Turbine Conference and Products
Show, Cleveland, Ohio, March 9-13, 1969, of The American Society of Mechanical Engineers.
Manuscript received at ASME Headquarters December 23, 1968.
Copies will be available until January 1, 1970.
THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS, UNITED ENGINEERING CENTER, 345 EAST 47th STREET, NEW YORK, N.Y. 10017
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The Engine Inlet on the 747
W. S. VIALL
INTRODUCTION
The design of an inlet with high pressure
recovery at low speed that can be enclosed in a
cowl with low external drag at high speed is the
subject of this paper. The inlet is an axisymmetric pitot type with a variable geometry auxiliary passage. It is installed on the high bypass
ratio front fan JT9D engine on the 747 airplane.
The nacelle inlet is integrated into the
total fan cowl and consists of the variable entry
lip, a duct to match the engine entry, and a low
drag external shape. Duct maximum Mach number,
duct diffusion, and surface pressure distributors
around the entry lip are considered and related to
auxiliary passage and diffuser design.
The variable entry lip consists of 12 pairs
of hinged doors located around the perimeter of
the cowl at its leading edge. During static and
low speed operation the doors are sucked open by
the engine permitting flow to enter through an
annular passage in addition to the main central
passage. Aerodynamic and mechanical aspects of
door actuation are discussed.
AERODYNAMIC CONSIDERATIONS
General
When an engine is selected for an airplane,
a great deal of effort is directed toward designing the most compact nacelle for it (2,8). 1 The
1
Numbers in parentheses designate References
at the end of the paper.
E
Fig.l Principal parts of the total fan cowl
front fan case is the largest portion of the engine that must be housed in cowling. Accessory
items that protrude may be faired locally and need
not affect the entire cowl diameter.
Subsonic inlet research and testing have established criteria for sizing a circular, pitottype inlet for a selected engine. This paper discusses inlets for subsonic engines cowled in strutmounted nacelles.
For low drag the desired inlet entry is
small in diameter. To save weight the cowl should
be minimal in length and simply constructed. The
engine demands a large airflow at full power.
Loss of pressure recovery during takeoff and inflight maneuvers that penalize airplane performance must be avoided. Minimum desirable pressure
recovery under cross-wind conditions at maximum
static takeoff power is 97 percent; pressure recovery must approach 100 percent below 100-knot
airplane velocity at takeoff power. Recovery
NOMENCLATURE
= highlight area
= throat area
A
TH
DHi = highlight diameter
DMAX = maximum cowl diameter
L\D = percentage increment of total airplane
drag
= actuator force
F
AQ
f(MTH ) = throat Mach number function
Fn = net thrust
M TH = throat Mach number
M oo = flight Mach .number
m = mass flow
m* = choking mass flow
ft4
P Top =
=
P T2
=
P
TH
Ps =
T =
Vu =
Voo =
WA =
✓
freestream total pressure
engine face total pressure
throat static pressure
wall static pressure
square root of total temperature
flow velocity at highlight plane
freestream or flight velocity
engine airflow
%AT= square root of temperature correction
to sea level standard day
= pressure correction to sea level
standard day
2
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TOP
INLET STALL
BOTTOM
O
C
ROTATION
ENGINE
OPERATES
WITHOUT
SURGE
0
A
HIGH RECOVERY - HIGH PERFORMANCE
A
4,8
300
350
Uz
150
200
250
300
350
AIRSPEED IKNOTSI
ENGINE
OPERATES
WITHOUT
SURGE
A
e,
100% RELATIVE
PRESSURE RECOVERY
0
0.90
0.90
FLIGHT MACH NUMBER
FLIGHT MACH NUMBER
Fig.2 Local inlet angle of attack
losses during climb and cruising flight are intolerable.
The 747 inlet is part of the total fan cowl,
as shown in Fig.l. The cowl consists of an entry
lip, a duct, and an external contour. Specific
parts of the fan cowl may be identified as follows :
1 Highlight--Location where lip leading
edge slope is normal to inlet centerline and fairs
into outer cowl.
2 Throat--Station where minimum duct flow
area occurs.
3 Diffuser--Internal duct between throat
and engine face.
4 Forebody--External cowl between highlight
and maximum cowl diameter station.
5 Boattail--External cowl between maximum
cowl diameter station and trailing edge of cowl.
6 Auxiliary Passage--Opening in cowl other
than main circular opening at front.
Although inlet design depends on many interrelated factors, the following is a logical sequence of consideration.
External Flow Field Definition
The environmental criteria by which inlet
performance is gaged should be established early
Fig.3 747 inlet performance envelope
in an inlet development program in order to assess
the risk associated in meeting recovery, distortion and blade stress requirements of the engine
over the entire airplane operating envelope.
The inlet flow field, in the broadest sense,
shall comprise a control volume which encompasses
the effects of the wing and strut (or pylon) in
the vicinity of the inlet; the upstream boundary
is in the undisturbed freestream and the downstream
boundary in the plane of the engine face. The
velocity and direction of flow at any and all
points in the defined control volume should be defined to provide information for answering the
following questions:
How does the flow demanded by the engine
approach and enter the inlet?
Having entered the inlet, how does the flow
approach and enter the fan?
What is the optimum orientation of the inlet in the flow field to give maximum recovery?
What is the impact of the inlet flow field
on nacelle dynamic stability?
During the development of the 747 inlets,
3
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1.01
1.00
At mint' = 0.995, 1% A W 4 - 8% A Mach
At mim• = 0.960, 1% A W x, - 2.5% A Mach
1.00
0.98
O
0.96
2 12
0.99
O
O
0.98
,AM,=
2.5% m
5
O
0.94
0.97
LT,
A
2
-4-
0.96
0.92
MTH
3%m
ASSUMED LIMIT-14
0
0.90
0.50
0.60
0.70
0.80
0.90
020
THROAT MACH NUMBER
Fig.4 Effect of orifice mass flow ratio on throat
Mach number
flow field analysis was conducted by means of computerized compressible axisymmetric flow programs,
wind tunnel model testing, and flight test of half
scale inlet cowls on the 707 airplane. Direction
sensing probes were used to survey the flow field
near the engine face and in the external flow regions.
Fig.2 shows curves pertaining specifically
to the 747 airplane which define local inlet inflow angle versus speed at an inboard nacelle.
These curves are representative of the specific
operating modes shown. There are other regions of
operation where conceivably inflow angles exceed
those shown. To establish criteria for a performance yardstick it was necessary to evaluate trial
configurations. Actual inlet-airplane demonstration flights were limited to finite conditions,
but meaningful model testing was accomplished in
advance of said demonstrations.
Fig.3 illustrates 747 operating bands. The
effects of pitch and yaw combined will produce
slightly higher resultant angles. The order of
magnitude of increase at stall is just over two
degrees, or more generally:
2
2
tan (yaw angle) + tan (pitch angle)
2
= tan (resultant angle)
The bands shown reflect this order of magnitude.
Engine operation in the regimes indicated
will vary for many circumstances; however, several
generalizations are pertinent. During takeoff and
climb, power settings are clearly defined, based
on engine manufacturer guarantees. During cruise,
power settings are based on range, payload, and so
forth, generally in the region of "max cruise."
On approach, engines are at or near idle. In general, when the engine is being worked the hardest
0.40
0.60
OBO
1.00
THROAT MACH NUMBER
1.00
Fig.5 Effect of throat Mach number on pressure
recovery
or in the region of low SFC, the highest achievable recovery is necessary. In emergency or otherwise unscheduled maneuvers when the airplane may
approach stall, it is necessary that the engine
continue to run without surge and produce sufficient thrust to aid in recovering to a normal
flight path.
The 747 inlet has been developed to provide
stall free engine operation and has met the engine
manufacturers requirements for fan blade stresses.
The inlet flow in no way produces nacelle dynamic
instability.
Throat Sizing
The engine airflow determines the throat
size. Engines in current use (1) and those being
offered for subsonic transport airplanes in the
1970 1 s are of the bypass type with a front-mounted
fan (2). Mach numbers at the fan range from 0.50
to 0.60. Lower bypass ratio engines (1.0 to 2.0)
require maximum mass flow at takeoff, while the
higher thrust demands at cruise of the high-bypassratio engines (3.0 and above) require maximum mass
flow in flight at Mach numbers from M O = 0.6 to
M O = 0.9 at cruise altitudes.
It is possible to match the throat of an inlet duct to any mass flow in the engine operating
envelope, but it is always required that the inlet
pass the maximum full-throttle mass flow obtainable in flight. The smallest throat passage that
will suffice is based on the throat Mach number
as defined by:
W VT
A
P - f (M
TH )
A
TH T
where A TM is the minimum geometric area of the
passage. M TH should not exceed 0.8.
Fig.4 shows the theoretical relationship of
mass flow ratio, m/m*, to the Mach number of the
4
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1.00
CAPTURED
STREAMTUBE
>
O
0.90
cc
HIGH-SPEED FLIGHT
TAKEOFF AND
LOW-SPEED FLIGHT
Fig.6 Captured streamtube comparison
a
E
0.80
30
flow passing through an orifice (3). The orifice
is choked; i.e., the Mach number approaches 1.0
as the value of mass flow approaches m*. At Mach
numbers approaching unity, small changes in mass
flow will produce successively larger changes in
Mach number; e.g., when m/m* = 0.96, a 1.0 percent airflow increase produces a 2.5 percent Mach num-
ber increase; while at m/m* = 0.99, a 1.0 percent airflow increase produces an 8.0 percent Mach number change,
The impact of throat Mach number on inlet
recovery, Fig.5, is an abrupt drop in performance
(3) just above Mach 0.8. Airflow is subject to increase on growth engines and may vary from engine to engine of a given rating by approximately ± 2 percent. It is important that the inlet throat be sized to pass the maximum airflow of high-tolerance engines in anticipation of reason-
able growth during the production run of the air-
plane, thus precluding an inlet redesign, The point in time when the 747 inlet throat
size had to be selected occurred before the engine
was sufficiently developed to target the airflow tolerance closer than ± 5 percent for the initial
rating. Airflow growth numbers were not clearly defined at that time. Full growth airflow was presumed, therefore, based on the given fan entry annulus area operating at a Mach number limit of 0.64, said limit being representative of current
compressor design technology. The subject of
throat sizing was treated from the standpoint of oversizing initially and getting engines with low airflow for first airplanes versus sizing to the
initially quoted flow and suffering penalties as
airflow increased with growth. Assuming equal
probability that the flow level of first engines
would be any value from -5 to +5 percent of ini-
tial nominal, and presuming that airflow will eventually increase beyond its initial value, an at-
tempt was made to evaluate the minimum risk incurred when compelled to size the inlet before the
exact airflow level is known. In no case would
the inlet throat be sized for less than nominal
airflow, so consideration was given to sizing for nominal or higher values,
40
50
INILETAMFLOW!'" (LB/SE
^141,5
Fig.7 Effect of lip area ratio on inlet pressure
recovery -- static free-stream conditions
Since inlet throat sizing is intimately tied
in with cowl lip sizing and they, together with
engine airflow variation, strongly affect cowl
drag, the impact of these parameters on minimum
risk to airplane performance can best be argued
from the standpoint of minimum cowl drag. On that
premise, quantitative analysis is discussed following the discussion of Lip Sizing philosophy.
Lip Sizing
The entry lip must accommodate a range of
inflow angles with no losses upstream of the
throat. The classic bellmouth inlet best meets
this requirement. It is simply a huge lip, and
the throat usually coincides with the compressor
face. Engine operation with a bellmouth, however,
is confined to static ground rig tests.
In high-speed flight, no appreciable losses
occur with a sharp entry lip upstream of a diffuser throat, but at low speeds some lip losses
are unavoidable. Fig.6 illustrates the two contrasting situations of captured streamtube. How
much lip loss can be tolerated depends on the
blade stress and surge margins of the selected
engine as well as the need for takeoff thrust.
If the throat size is maintained and the lip made
blunt, the highlight diameter approaches the minimum diameter possible to cowl the engine. A high
ratio of highlight diameter to cowl diameter
causes high drag at cruise Mach numbers. Depending on the design cruise Mach number of the airplane in question, the designer may choose to
favor either a fat lip for low-speed or a thin
lip for high-speed operation. In the latter case,
some auxiliary inlet device will be required for
takeoff and approach.
For the static case, the loss of inlet pressure recovery as engine airflow increases and
/A
"fixed-lip" area ratiodecreases
A hilight throat
is shown in Fig.7 (3). The average total pressure
5
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▪
M m = 0.90 at 35,000 feet
•
•
Tolerance
Li
Growth
0
1,760
Nominal
M%lip isbasehnec onfiguration
1,740
ALTITUDE I
M TH = 0.826 3 ,
3, = Total fan cowl drag with throat sized to accommodate nominal climb airflow
1,720
with engine operating at nominal cruise airflow .
O
= Penalty for oyersizing throat to accommodate greatest climb airflow with
O
engine operating at least cruise airflow.
0
0
1,700
= Penalty for throat sized to accommodate nominal climb airflow with engine
operating at greatest cruise airflow.
0'
trading margin
A0
between 9% and 18% lip
1.680
rf
A
O
0
O
ALTaUDEMMOFEET/
1,660
2
40
35
C
0
1,640
;;
30
1,620
NOMINAL
9/ LIP
9% LIP
18% LIP
18% LIP
Fig.8 Engine airflow tolerance and growth effect
on airplane drag
1,600
1,580
040
0.50
0.60
0.70
0.80
0.90
1.00
FLIGHT MACH NUMBER
Fig.l0 JT9D corrected airflow at cruise rating
0
AIRFLOW-5
A
0
A
ZERO REF
FOR NOMINAL
THROAT
`7`t
7'f
0
2
CS
2
Common
D ina%
Constant mach number
INLET TYPE
•
18% lip Blow-in door
Constant altitude
Growth sized throat
■
9% hp Slotted lip
2.15%-oyersized throat
•
5% lip Slotted lip
Nominal throat
Fig.9 Carpet plot relating inlet drag to airflow
and lip size
recovery is measured at the compressor face and,
therefore, includes both lip and diffuser losses.
The auxiliary inlet offers a means of avoiding
low-speed recovery losses of thin-lipped cowls
and will be discussed in the following section.
Cowl Sizing Trade Studies
In exercising the various fan cowl sizing
trades for the JT9D short duct installation, we're
talking about total fan cowl drag of from 3.5 to
5.0 percent of airplane drag at M oo = 0.9 and
35,000 ft with engines operating at MCR rating.
An additional inlet drag penalty to the
first airplane would exist due to the early 5
percent airflow tolerance quoted. To meet takeoff and cross-wind requirements, the "707-type"
18 percent lip with a blow-in door auxiliary in-
let was considered as the baseline. With future
airflow growth, a penalty of from 0.5 to 1.0 percent in addition to that for airflow tolerance
could exist.
Assuming that the worst case for an 18 percent lip inlet has occurred, then the initial risk
is equal (in terms of drag penalty) whether the
throat is sized for the +5 percent level or sized
at nominal (Fig.8). However, in the latter instance, any further airflow growth results in more
drag; while, if the throat is oversized, an airflow increase reduces the penalty from that suffered initially. Choosing a thinner lip inlet
design to favor low drag, an oversized throat with
a 9 percent lip has the advantage initially as
well as thereafter over an 18 percent lip inlet.
Fig.9 is a carpet plot which relates inlet parameters of throat size, lip ratio, and airflow variation to a drag parameter.
The ordinate of the plot, AD, is change in
lip suction plus super-velocity (L.L.S. + S.V.)
drag (4,6,7). This is only a portion of total fan
cowl drag, but it is that portion most strongly
influenced by inlet lip and throat sizing.
Several qualifying assumptions have been
made in Fig.9. These are:
1 Constant airplane Mach no.
2 Constant altitude
3 Cruise thrust rating
4 Fixed fan cowl maximum diameter
5 The zero reference line is the 18 percent
lip blow-in door inlet drag level at nominal engine airflow and with a nominal throat size.
6
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FOREBODY 1,760
•
BOAT TAIL
AUXILIARY
PASSAGE
1,740
...
1,720
ii
HIGHLIGHT
1,700
1,680
1,660
1,640
1.00
1,620
AUXILIARY
PASSAGE
INCREMENT
FOR 1.18 LIP
NOMINAL
1,600
0 90
1,580
0,40
0,50
0.60
0.70
ORO
090
1.00
FLIGHT MACH NUMBER
Fig.11
JT9D corrected airflow at climb rating
0 80
30
40
Wq INLET AIRFLOW, AUM)SEC.FT 2I
Fig.9 illustrates varying captured mass flow
inlets. Point A is a drag level with the reality
of having to oversize the throat of the "standard"
inlet to accommodate a possible +5 percent airflow
but with the actual airflow .having turned out to
be - 5 percent.
There are three "carpets" on the plot, each
associated with a throat size and several values
of airflow. At the left edge of each carpet,
large dark symbols indicate three lip area ratios.
Moving from left to right on a given carpet, the
right-hand edge is determined by the throat Mach
number limitation as airflow increases.
Particular inlet configurations which allow
drag reductions are indicated by Points B and C
Fig.13 Effect of lip area ratio on inlet pressure
recovery static free-stream conditions
Variable Geometry Devices
The 707-type inlet (3) with an auxiliary
passage aft of the diffuser consisting of a ram
fed, converging passage, which is closed during
cruise by aerodynamically actuated doors, is shown
in Fig.12. During low-speed operation, airflow
bypasses the main lip, reducing lip and diffuser
losses as shown by the shaded area in Fig.13.
This device requires an 18 percent or larger lip
area ratio (2) in order to achieve acceptable engine performance.
and are described as follows:
To offset the cruise drag penalty due to
oversizing, the lip area ratio must be reduced at
(B) slotted lip inlet with a 9 percent lip
least 1.09 (reducing the highlight diameter),
(C) slotted lip inlet with a 5 percent lip.
whereupon it becomes necessary to adopt a new
Figs.10 and 11 show plots of engine corrected means to reduce low-speed lip losses. A 9 percent
airflow versus flight Mach number at various alti- lip on the 707 blow-in door inlet would penalize
take-off thrust in the order of 10 percent.
tudes for the MCR and MCT ratings, respectively, Studies were made of means to reduce lowIn the instance of inlets with the throat over-
speed lip losses by using boundary layer control
sized, the corrected airflow, which produces the
(suction and blowing) and other devices such as
design limit throat Mach number of 0.7j, is indiinflatable lips. The concept of a lip slot, ancated by a horizontal line. Airflows of both
alogous to a wing leading edge slot, had been
nominal and 5 percent in excess of nominal are
tested oxtensively as to low-speed performance
shown. The regions on each plot above the 0.77
throat Mach number line represent risk areas where with impressive results. Both wind tunnel model
and prototype flight testing demonstrated the
choking losses inside the inlet will result in enfeasibility of such a device. The 747 inlet congine thrust loss; thus, a risk-of-retooling area
figuration was selected based on this concept.
could be defined.
7
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Fig.14 Prototype fixed-leading-edge inlet
Fig.15 Prototype rotating-leading-edge inlet
Two prototypes are illustrated in Figs.14 and 15. cussion maps out the environmental conditions
In principle, the slotted lip inlet embodies under which the doors should open and the general
12 sets of inner and outer doors pivoted at the level of working mediums which cause the doors to
cowl leading edge in such a manner that the entire move.
cowl entry lip (except for support struts) rotates
For the purpose of discussion, let us assume
outward petal fashion, as the doors assume the
that all doors are held closed initially by a hyopen position. The actual 747 configuration has
pothetical actuator force of 500 lb and that an
a fixed leading edge ring which allows a more ruginput is made which should cause them to open.
ged and reliable design. Comparative section views
Consider the working media which will move
are shown in Fig.16. Engine performance with this
the doors; a force holding them closed which reconfiguration is down slightly compared to the rosults from deflection of a spring by means of a
tating lip type, and greater mechanical force is
system of mechanical levers and a force tending
required to close this type of door at the desired
to pull them open resulting from a pressure drop
Mach number.
(AP) acting on the "flat plate" area of the doors.
The performance of both types was evaluated
The directional sense of the opening "AP"
by means of large-scale wind tunnel and flight
results from low static pressure on the engine
test models. Internal pressure recovery and movside of the doors relative to that on the external
able door actuation were measured and observed.
side of the doors. Local velocities around the
Results showed that the concept will produce
doors not withstanding, door opening force potenexcellent pressure recovery throughout the taketial develops as a function of the relationship
off, climb, and cruise modes. One qualifying conbetween P oo and P th , i.e., ambient static pressure
dition exists. It is desirable that the doors go
versus throat static pressure. Local velocity efshut during climb at an airplane speed of Mach 0.5
fects require more complicated analysis than is
Testing showed that the doors will close at Mach
argued at this time.
0.56 if the inflow is axisymmetric. If the inflow
Fig.17 is presented in order to illustrate
angle is not zero, the "upwind" doors will remain
that the door opening medium deteriorates with inopen, while doors on the "downwind" side will
creasing altitudes, while the mechanical closing
close.
medium is independent of altitude. Consider the
Assessment of 747 inflow angles during climb
curve labeled "takeoff." The opening force funcshows that the nacelles would see upwash of two to
tion exceeds the mechanical closing function at
four degrees. Accordingly, in order to shut the
takeoff power setting from 0 to 10,000 ft altitude;
lower doors of the inlets, mechanical means
i.e., the doors are potentially open. The curve
(springs, and so forth) had to be employed.
labeled climb at Mach 0.4 in. shows how, if climb
Considerable effort was devoted to inlet lip
power is selected while the airplane is flying at
door closure during climbout following takeoff,
Mach = 0.4, the opening medium decreases with alas well as the need for subsequent door opening.
titude. We shall presume that the door opening
In order to illustrate the type of problems faced
requirement is that velocity ratio (V 1 /V 0p ) be
in designing a workable inlet, the following disgreater than unity in which case, when flying at
8
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El Approach envelope
Holding envelope
ROTATING
LEADING EDGE
---- Domestic climb schedule below 10,000 feet
ALINES OF VELOCITY
RATIO EQUAL
TO UNITY FOR'
50% CRUISE
POWER
Foreign climb schedule
— — Constant-airspeed climb
CRUISE
POWER
25
20
CLIMB
POWER
TAKEOFF/
POWER
/
VELOCITY
RATIOS
GREATER THAN
UNITY
J
/
VELOCITY
FIXED
LEADING EDGE
INNER - DOOR PIVOT
Fig.16 Schematic comparison of fixed and rotating
leading edges
FLIGHT MACH NUMBER
Fig.18 Relationship of unity velocity ratio to
747 flight envelope
HYPOTHETICAL
DOOR ACTUATING
FORCE (LB)
UO
UZ
CLIMB POWER.
AIRSPEED Mn= 0.7
EQUILIBRIUM
EXAMPLE 500
to )
a
20
0
0
rc
TAKEOFF POWER,
= 0.4
AIRSPEED M
O
TAKEOFF POWER,
STATIC
10
20
CLIMB POWER,
AIRSPEEDS THAT
PRODUCE VI =
10
V,
CLIMB POWER,
AIRSPEED M, = 0.4
P rN
TAKEOFF
30
ALTITUDE 11.000 FT)
Fig.17 Deterioration of door opening force with
altitude
Mach = 0., there is no requirement for the doors
to open. On the other hand, when flying at Mach
= 0.4, selection of climb power will always require that the doors be open. Whether or not they
do, in fact, open is then dependent upon the balance of mechanical closing and pressure opening
media. For our hypothetical example of a mechanical closing medium of 500 lb, the doors would not
open above 17,000 ft altitude.
Having stipulated that V i /V oo greater than
unity is the door opening criteria, it is pertinent to map the relationship of unity velocity
ratio to the 747 flight envelope. This is the
intent of Fig.18. The solid, near-vertical lines
represent unity velocity ratio for various power
settings. Regions of holding and approach and
two takeoff schedules are also shown. If the aircraft is operating at a point on the map to the
left of a particular "unity line," the velocity
ratio is greater than unity for that particular
Mx=
'TN
CLIMB
0.0
0.4
ALTITUDE 11,000 FEET)
Fig.19 Absolute pressure levels and equivalent
spring actuator forces versus altitude
power. Operation to the right produces less than
unity velocity ratios. For example, the aircraft
is holding at point "j" -- 23,000 ft at Mach 0.45;
if the cruise power setting were of the order 50
percent cruise, then "j" is to the right of the
unity line (less than unity) and the doors should
be shut. If, now, climb power is selected, point
"j" is to the left of that unity line and the
doors should open. Referring to Fig.17, we see
that if the actuator force is 500 lb the balance
media will be such as to prevent door opening.
Returning to Fig.17 and supposing an approach condition at point "k," doors would be shut
for a power of 50 percent cruise but should takeoff power be selected, the doors should open.
9
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INLET
o
FLAPS
DOWN
d
FLAPS UP ACCELERATE & CLIMB
.1
CRUISE
STAN
CONTH
V2 CLIMB
V7
150
200
om am
300
250
Mx
AIRSPEED (KNOTS)
0.70 0.80
am
FLIGHT MACH NUMBER
Fig.20 Inlet angle of attack at 747 climb schedule
00
0.40
050
060
FLIGHT MACH NUMBER
0.70
Fig.22 Net moments at 747 climb schedule—bottom
doors only
DESIRED DOORS-OPEN REGION
O
747 CLIMB SCHEDULE
2
t`>
'
Opening: P
DOORS
THEORETICAL 1_
: OPEN
.
8
o
0
040
0.50
0.60
0.70
- P
th
- P
Closing; F
act
co
DOORS
CLOSED
080
FLIGHT MACH NUMBER
Fig.21 Inlet velocity ratio at JT9D climb schedule
Referral to Fig.18 shows the balance of media is
in favor of doors open.
The two constant speed climb schedules,
shown in Fig.18, represent the domestic and for► mode of operation, respectively.
eign operators
A major difference is seen in that, during the
former, the velocity ratio will be greater than
unity until 10,000 ft altitude and 0.48 Mach is
reached; while, in the latter case, the crossover
occurs at approximately 1200 ft and 0.49 Mach.
The working medium balance is in favor of doors
closed at both crossover cases. See points 11p"
and "q," respectively.
Fig.19 illustrates the level of absolute
pressures for various altitudes and Mach numbers.
freestream static
Shown are throat static (th)'
P
'
(Poo ) and actuator force (Fac t).
are labeled either "Takeoff"
Lines of P
th
or "MCT" because they represent the isentropic
0.2M2]3.5 for the airflow exfunction PTLl
istent at those power settings. The P oa line is
from standard atmospheric tables. The
"Pact" line
is computed by means of converting the single
point actuator force acting on the inner door to
an equivalent equally distributed pressure drop.
The working media discussed in the foregoing are
as follows:
oo
The assumptions made in the foregoing discussion
are purposely oversimplified in order to bring
the whole spectrum of door operation into focus.
Many factors contribute to the actual inlet-dooron-the-airplane forcing functions including local
velocities (already mentioned) angle of attack,
mechanical friction, and so forth.
To obtain actual door actuation force data,
wind tunnel and flight testing was conducted using
models and prototype hardware with movable doors
which were instrumented with surface pressure taps
as well as strain gages. Hinge moment calculations were done based on these data and from them
the actuator (spring) force determined that it is
necessary to close the doors. Each pair of inner
and outer doors is linked together such that both
doors reach the ends of their respective (unequal)
excursions simultaneously. The actuator force is
input at a convenient point on the linkage.
Fig.20 shows the unrestricted speed-climbout
schedule (foreign operation) versus inlet angle of
attack. Fig.21 shows the inlet velocity ratio
schedule for that climbout. Figs.22 and 23 are
the calculated door pair hinge moments for bottom
and top doors, respectively. It is readily apparent that the top doors will close at a lower Mach
number than the bottom doors; furthermore, in
order to close the lower doors, the actuator force
must increase as the doors move from open to
closed. The reverse is true of upper doors.
Diffuser Design
One means of defining the shape of the in-
10
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8
1.05
9.:
I///
20
0
60
40
80
100 110
PERCENT OF DUCT LENGTH
Fig.24 Variation in diffuser wall static pressure
0 40
0 50
0.60
0 /0
BOTH INLETS
(DOORS CLOSEDI-
FLIGHT MACH NUMBER
Fig.23 Net moments at 747 climb schedule-top
doors only
\
1.00
0.99
ternal diffuser from the throat to the engine is
discussed in NACA TN-3668 (4). Adjustments in the
diffuser exit are necessary to account for Lhe
area of the spinner on the first fan stage, since
NACA TN 3668 treats diffusers without center
bodies.
Diffuser area ratios of up to 1.25 are
reasonable when following NACA TN-3668. For
greater diffusion, other design techniques may be
explored.
Duct area change should begin gradually behind the throat for good diffuser performance at
low forward speed. It is important that the wall
slope change be gradual behind the throat. Once
the desired diffusion angle is reached, the wall
can become conical before fairing into the engine
face. A conical section is desirable from a
structural standpoint. Fig.24 shows wall static
pressure data measured in a full-scale inlet diffuser contoured according to the foregoing criteria (3). The uniformly rising static pressure
level along the length of the duct indicates satisfactory diffusion.
0.98
Fr'
0.97
0.96
-
Inlet Total Pressure Recovery
Fig.25 illustrates wind tunnel test data on
engine face pressure recovery during takeoff, rotation, and climb-out for the two configurations
studied. Both airflow and angle of attack schedules are implicit in this plot. From the intersections of the "doors open" and "doors closed"
curves, the desired door closing Mach number is
determined for each configuration, i.e. 0.36 and
0.40. Failure to close the doors at those speeds
results in intolerable recovery losses.
FIXED LEADING-EDGE
INLET (DOORS OPEN)
ROTATING LEADING-EDGE
INLET (DOORS OPEN)
0.95
0
0.10
0.20
0.30
0.40
0.50
0. 60
03 0
FLIGHT MACH NUMBER
Fig.25 Pressure recovery comparison of two inlet
types
The final 747 production inlet configuration
is designed so that the doors will close during
climb-out and provide best possible recovery. At
those times in the flight envelope when the doors
are normally closed but step changes require them
to open, they will do so.
REFERENCES
1 Viall, W. S., "Development of a Subsonic
Diffusing Engine Inlet for the 707-120B/720P Airplanes," Document D6-7588, The Boeing Company,
Sept. 10, 1962.
2 Viall, W. S., "Aerodynamic Considerations
for Engine Inlet Design for Subsonic High Bypass
Fan Engines," presented at SAE meeting, Los Angeles,
Oct. 1966.
3 Klees, G. W., "Subsonic Inlet Research
and Development-Variable Concepts Applicable to
707 Airplanes with Advanced JT3D Fan Engines,"
Document D6-8983, The Boeing Company, Oct. 15,
1964.
4 Scherrer, R. and Anderson, W. E., "Pre11
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liminary Investigation of a Family of Diffusers
Boeing Company, Jan. 5, 1966.
Designed for Near Sonci Inlet Velocities," NACA
7 Monk, J. R., "Estimation of Subsonic EnTN-3668, Feb. 1956.
gine Nacelle Drag," D6-8057, The Boeing Company,
5 Schlicting, H., "Boundary Layer Theory,"
Sept. 5, 1963.
McGraw Hill, New York, 1960, Chapters XXII and
8 Frazier, G. T., "Aerodynamic ConsideraXXIV.
tions for Engine Exhaust Design for High-Bypass
6 Lawrence, R. L., "Nacelle Cowling of High- Subsonic Fan Engines," presented at SAE meeting,
Bypass-Ratio Turbofan Engines," D6-18086TN, The
Los Angeles, Oct. 1966.
12
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