Aircraft cabin pressure control for ascents and descents

The rate of change of cabin air pressure in an aircraft cabin during aircraft ascent and descent is controlled as a function of a ratio which is equal to (P.sub.ld -P.sub.c)/(P.sub.ld -P.sub.a) during descent, and is equal to (P.sub.cc -P.sub.c)/(P.sub.cr -P.sub.a) during ascent, where P.sub.ld is ambient pressure at the aircraft landing site, P.sub.c is cabin pressure, P.sub.a is external aircraft ambient pressure, P.sub.cc is cabin pressure at cruise altitude, and P.sub.cr is ambient pressure at cruise altitude. A set point value for the ratio is computed throughout aircraft ascent and descent, and compared to the ratio to schedule a variable flow area of an outflow valve in order to drive the difference (i.e., the error) between the ratio and the desired value for the ratio to zero. By using an adjustable value for the desired set point, control of cabin pressure rate of change is improved during ascent and descent, since the cabin pressure rate of change is maintained at a reduced, linear rate, thereby providing a more comfortable environment for passengers.

TECHNICAL FIELD 
This invention relates to controlling aircraft cabin pressure and more 
particularly to computing an adjustable set point reference for the cabin 
pressure rate of change during aircraft ascents and descents. 
BACKGROUND ART 
Air pressure within an aircraft cabin is controlled during ascents and 
descents as a function of: ambient pressure, aircraft cabin pressure and 
the pressure at the aircraft landing site. Control is typically performed 
by an automatic cabin pressure controller. 
As described in U.S. Pat. No. 3,473,460 to Emmons, and assigned to the 
assignee of the present invention, an aircraft is designed to withstand a 
maximum pressure differential between the cabin and ambient pressures. The 
'460 patent discloses a system where the desired rate of cabin pressure 
change, d(P.sub.cd)/dt is computed as a function of a ratio DP.sub.c 
/DP.sub.a, where: 
DP.sub.c =the remaining change in cabin pressure during either an ascent or 
descent; and 
DP.sub.a =the remaining change in ambient pressure during either an ascent 
or descent; 
The '460 patent discloses controlling the rate of cabin pressure change and 
adjusting the rate so if the aircraft slows its ascent/descent rate, the 
rate of cabin pressure change will also slow. Similarly if the aircraft 
increases its ascent/descent rate, the rate of cabin pressure change will 
also increase. This ensures the aircraft fuselage does not exceed a 
predetermined maximum pressure differential between the cabin pressure and 
external ambient pressure. 
However, the '460 patent uses a function generator (item 33, FIG. 1) having 
a single, nonadjustable operating line as a function of DP.sub.a to 
provide a set point for the desired rate of cabin pressure change, 
d(P.sub.cd)/dt. Setting the desired rate of cabin pressure change based 
upon this single operating line for an infinite number of possible cabin 
pressure ascent and descent profiles, leads to a sacrifice in the control 
of cabin pressure since the desired rate of cabin pressure change is 
constrained to take on values along the single nonadjustable operating 
line of the function generator. 
DISCLOSURE OF THE INVENTION 
An object of the present invention is to provide an aircraft cabin pressure 
controller having an adjustable set point reference value for the cabin 
pressure rate of change during aircraft ascent and descent. 
According to the present invention, the rate of change of aircraft cabin 
pressure during aircraft ascent and descent is controlled as a function of 
the difference between (1) a ratio which equal to the remaining change in 
cabin pressure divided by the remaining change in ambient pressure (i.e., 
DP.sub.c /DP.sub.a), and (2) a set point reference value which is a 
function of the remaining change of ambient pressure and the initial value 
of the remaining change in cabin pressure at the start of the ascent or 
descent; the difference value is summed with a constant value indicative 
of the desired initial cabin pressure rate of change to provide an 
adjustable set point reference value, the adjustable set point reference 
value is a set point for the actual cabin pressure rate of change and is 
computed at successive intervals through out the ascent and descent. 
Using an adjustable reference as a set point for controlling the cabin 
pressure rate of change produces a more comfortable environment for 
passengers, since cabin pressure is changed at a reduced, more constant, 
linear rate during aircraft ascent and descent. 
These and other objects, features and advantages of the present invention 
will become more apparent in light of the following detailed description 
of a best mode embodiment thereof as illustrated in the accompanying 
drawings.

BEST MODE FOR CARRYING OUT THE INVENTION 
FIG. 1 is a functional block diagram an aircraft cabin pressure system 10 
according to the present invention. The system 10 includes an electronic 
cabin pressure controller 12 (hereinafter "controller") and an outflow 
valve 13. The controller 12 includes: a microprocessor 14 (e.g., Intel 
80186, or Motorola 68020), memory 16 (e.g., PROM, RAM), and input/output 
ports 18,20 which include analog-to-digital and digital-to-analog 
converters. The elements are interconnected through an address/data bus 
21. The controller provides an output command signal on a line 22 to the 
outflow valve 13 which has a variable flow area through which pressurized 
cabin air is discharged to regulate actual cabin pressure. An example of 
such a valve is disclosed in U.S. Pat. No. 3,740,006 to Maher, and 
assigned to the assignee of the present invention. 
Throughout this disclosure various signals will be continuously referred 
to, and in an effort to facilitate the disclosure the following 
nomenclature information is provided: 
______________________________________ 
NOMENCLATURE TABLE 
VARIABLE EXPLANATION 
______________________________________ 
C.sub.f Correction factor 
d(P.sub.cd)/dt 
Adjustable set point reference 
for the cabin pressure rate 
of change. 
ECPC Electronic Cabin Pressure 
Controller (i.e., the 
controller) 
M Ascent/descent starting point 
factor 
P.sub.a External aircraft ambient 
pressure 
P.sub.c Cabin pressure 
P.sub.cr Ambient pressure at cruise 
altitude 
P.sub.cc Cabin Pressure at cruise altitude 
P.sub.cd Desired cabin pressure 
P.sub.ld Ambient pressure at the aircraft 
landing site 
P.sub.r Desired value of the ratio 
DP.sub.c /DP.sub.a 
DP.sub.a = (P.sub.1d - P.sub.a) during a descent 
= (P.sub.cr - P.sub.a) during an ascent 
DP.sub.c = (P.sub.1d - P.sub.c) during a descent 
= (P.sub.cc - P.sub.c) during an ascent 
______________________________________ 
The controller 12 receives aircraft data from aircraft sensors and avionics 
24; the data includes: ambient pressure at the aircraft cruise altitude 
(P.sub.cr) on a line 25, ambient pressure at the aircraft landing site 
(P.sub.ld) on a line 26, cabin pressure at cruise altitude (P.sub.cc) on a 
line 27, sensed ambient pressure (P.sub.a) on a line 28, sensed cabin 
pressure (P.sub.c) on a line 30, and a signal on a line 31 indicative of 
whether the aircraft ascent/descent is complete. These six signals may be 
provided over the aircraft's digital data bus (e.g., ARINC 429, ARINC 629 
or MIL-STD-1553) (not shown) if so equipped, or from dedicated electrical 
lines presented to the controller, or in any other appropriate manner. 
The present invention is best disclosed by first functionally describing 
the desired operation of the system, and then illustrating an embodiment. 
Therefore a functional description of the present invention will be now be 
provided. 
The controller 12 acts to maintain the magnitude of the cabin pressure rate 
of change at a value less than an adjustable set point reference. The 
expression for the adjustable set point reference is: 
EQU d(P.sub.cd)/dt=C.sub.1 +C.sub.f (Eq. 1) 
where 
##EQU1## 
The correction factor, C.sub.f is defined as: 
EQU C.sub.f =K.sub.1 *((DP.sub.c /DP.sub.a)-P.sub.r) (Eq. 2) 
where: 
##EQU2## 
When the ratio (DP.sub.c /DP.sub.a) deviates from the set point value 
P.sub.r, C.sub.f becomes non-zero to add the necessary correction factor 
in order to drive the difference between the ratio (DP.sub.c /DP.sub.a) 
and P.sub.r towards zero. 
P.sub.r is derived from aircraft ascent and descent profiles, and has a 
different equation for each of the two phases. Furthermore, each aircraft 
type may have different equations for P.sub.r since each type typically 
ascends/descends in a slightly different manner. As an example, the long 
haul Boeing 747 series aircraft certified to operate up to approximately 
45,000 feet, descends in a different manner than the short haul Boeing 737 
series aircraft which is certified up to approximately 35,000 feet. Both 
in turn ascent/descend in a different manner than a Short Takeoff and 
Landing (STOL) aircraft such as the DASH-7 manufactured by the De 
Havilland Corp. Therefore, based upon the typical descent profile for a 
particular aircraft type, an equation can be derived for P.sub.r. However, 
it should be understood it is possible to use one standard equation for 
all large aircraft, but the trade-off is the equation may not have the 
accuracy you desire from the controller since the equation covers such a 
broad range of aircraft. Therefore, embodiments of the present invention 
use different equations to define P.sub.r during ascent and descent for 
each aircraft type (e.g., Boeing 777, Airbus 340, etc.). 
FIG. 3 is an illustration 34 of a typical aircraft flight profile from 
which equations for P.sub.r during ascent and decent can be derived. 
Altitude 36 and ambient pressure (i.e., P.sub.a) 38 are plotted along the 
vertical axes versus time 40 on the horizontal axis. The aircraft flies 
along a curve 42 which represents altitude and ambient pressure versus 
aircraft flight time. The aircraft takes off at a point 44 and climbs 
along the curve 42 to a cruise altitude 46 and levels off. The aircraft 
then enters the cruise phase of the flight flying at substantially 
constant altitude. When the aircraft reaches a point 48 it starts the 
descent phase of the flight. The aircraft descends until it levels off at 
an altitude 50 for a short while, usually due to air traffic control 
instructions, and then starts the descent again until the aircraft lands 
at a landing field point 52. The pressure at the landing field is 
P.sub.ld. 
During the flight P.sub.c is changing along a curve 54 under the control of 
the controller 12 as a function of P.sub.a. At the aircraft maximum 
operating altitude of 45,000 feet, P.sub.c reaches a maximum of 8,000 
feet. When the aircraft starts to descend from its cruising altitude of 
45,000 feet, the cabin starts to descend from its pressurized value of 
8,000 feet. 
The cabin typically lands at a point 56 before the aircraft landing point 
52. This ensures P.sub.c is at a slightly higher pressure than P.sub.a 
when the aircraft lands, thereby allowing the aircraft doors to be opened 
easier in case of an emergency. 
During the non-linear ascent and descent of the aircraft along the curve 
42, it is desirable that the rate of change of P.sub.c with respect to 
time be constant (i.e., C.sub.f =0). This is particularly true during the 
descent phase where the human ear is more sensitive to pressure changes, 
and hence a reduced, linear, cabin pressure rate of change leads to an 
increase in passenger comfort. An exemplary constant cabin pressure rate 
of change during a descent is approximately -300 feet per minute (fpm). 
While an exemplary constant cabin pressure rate of change during an ascent 
is approximately 500 fpm. 
The controller limits the rate of change of P.sub.c during ascent and 
descent by monitoring DP.sub.c and DP.sub.a. As mentioned hereinbefore the 
general form of the equation equated to P.sub.r is determined based upon 
the typical aircraft descent or ascent profile, and it relates the 
nonlinearity of the actual aircraft ascent/descent to the desired linear 
ascent/descent of the aircraft cabin. 
Referring still to FIG. 3, during the descent starting at the point 48, the 
ratio DP.sub.c /DP.sub.a, can be plotted as a function of DP.sub.a. The 
result of such a plot is FIG. 4. FIG. 4 is a plot 58 of P.sub.r along the 
vertical axis 60, versus [P.sub.ld -P.sub.a ] (i.e., descent DP.sub.a ) 
along the horizontal axis, 62. While operating at 45,000 feet with an 
8,000 foot cabin, the aircraft starts its descent from a point A 64 and 
follows a first descent curve 66 until reaching a point C 68, which 
represents the landing field. At point C 68, the aircraft is on the ground 
so DP.sub.a is equal to zero. The curve 66 was derived by plotting the 
values of DP.sub.c /DP.sub.a from FIG. 3 as a function of DP.sub.a. As an 
example referring back to FIG. 3, the value of DP.sub.c /DP.sub.a at the 
point 48 equals 0.30 psi/psi which corresponds to the value of DP.sub.a 
equal 12.56 psi. These values define the point A 64 of FIG. 4. Repeating 
this procedure for a plurality of points along the curve 42 (FIG. 3) 
during the descent phase of the flight defines the curve 66 in FIG. 4. 
If the aircraft starts its descent from an altitude of less than its 
maximum operating altitude of 45,000 feet, then P.sub.r is defined by a 
curve other than the first descent curve 66. In general, there are a 
plurality of curves (not shown) each starting from a locus 70 of possible 
descent starting points and each ending at a particular landing field 
point (e.g., point C 68). An example of one of these plurality of curves 
is a curve 72 which represents a descent starting from a point B 74 to the 
landing field point C 68. Point B 74 represents a situation where the 
aircraft fuselage has a maximum differential pressure across it while 
operating at an altitude less than the aircraft's maximum operating 
altitude. For situations where the aircraft fuselage is not experiencing a 
maximum pressure differential, a locus of possible starting points is a 
second locus 76. 
Another possible locus of descent starting points is defined by a locus 
running from the point A to a point E 78. This locus represents the 
situation in which the aircraft cabin is at 8,000 feet while the aircraft 
is at an altitude below its maximum operating altitude (e.g., 45,000 
feet). 
Once a curve for P.sub.r is known by plotting a sufficient number of 
points, an equation for P.sub.r can be derived which defines the family of 
curves originating from a point along the locus of starting points (e.g., 
70 and 76) and terminating at the landing field (e.g., point C 68). The 
equation for P.sub.r can be derived using known curve fitting techniques, 
or one may empirically derive an equation for P.sub.r based on the curve 
66. In general the higher order the equation for P.sub.r, the less error 
there will be in the curve fit equation. In this embodiment a second order 
equation for P.sub.r is used, which is defined as: 
EQU P.sub.r =[C.sub.2 -(M*DP.sub.a)+K.sub.2 *(DP.sub.a -C.sub.3).sup.2 ](Eq. 3) 
where: 
C.sub.2 =unitless 
C.sub.3 =psi 
K.sub.2 =1/(psi).sup.2 
M=ascent/descent starting point factor. 
Note: C.sub.2, C.sub.3 and K.sub.2 are constant values. 
This equation characterizes curves 66 and 72, and the plurality of curves 
(not shown) indicative of a reduced, linear, cabin pressure rate of change 
during a descent. 
A line 79 tangential to a portion of the curve 66 is characterized by the 
equation for M. M (the ascent/descent starting point factor) is the 
coefficient for the term raised to the first power of DP.sub.a in Eq. 3. 
The line 79 varies based upon the starting point of the descent (e.g., A 
64) and the end point of the descent (e.g., C 68). For a descent, M is a 
function of the descent starting point (e.g.,point A 64) from the locus of 
possible starting points. Therefore M is only computed once, at the start 
of the descent, and its value is retained until the descent is completed. 
Similarly for an ascent, M is computed once at the start of the ascent as 
a function of the ascent starting point. M provides "memory" in Eq. 3 of 
the ascent or descent starting point. The equation for M is derived by 
substituting Eq. 3 into Eq. 2 and solving for M when C.sub.f is set equal 
to zero, which yields: 
EQU M=[C.sub.2 +K.sub.2 *(DP.sub.a -C.sub.3).sup.2 -K.sub.1 *(DP.sub.c 
/DP.sub.a)]/DP.sub.a (Eq. 4) 
With respect to an ascent, FIG. 5 illustrates a plot 80 of the ratio 
DP.sub.c /DP.sub.a along a vertical axis 82, and DP.sub.a along a 
horizontal axis 84 during the ascent phase of the aircraft flight along 
the curves 42, 54 (FIG. 3). Plotting the ratio DP.sub.c /DP.sub.a as a 
function of DP.sub.a as the aircraft ascends from a point 44 (FIG. 3) to 
the cruising altitude 46 results in a curve 86. The curve 86 is just one 
example of an ascent curve, and as such it should be understood that there 
are a family of ascent curves based upon the starting and ending points of 
the aircraft ascent. An explanation will now be presented of how the 
present invention computes a rate limit set point for the cabin pressure 
rate of change. 
In FIG. 2 is illustrated logic 32 stored in the memory 16 and executed by 
the microprocessor 14 to control the cabin pressure rate of change during 
aircraft descents and ascents. Upon entering the logic 32 at a step 90, 
the ascent/descent starting rate C.sub.1 is set to an initial value for 
the limit of the rate of cabin pressure change (e.g., -300 ft/min for 
descent, and 500 ft/min for ascent) in a subroutine 92. A subroutine 94 is 
then performed to read the five pressure signals: P.sub.a, P.sub.c, 
P.sub.cr, P.sub.cc and P.sub.ld. Next a subroutine 96 is performed to 
initialize the correction factor C.sub.f to zero by computing the 
ascent/descent starting point factor, M at the start of the ascent or 
descent. Subroutine 96 computes M as defined in Eq. 4 for the particular 
DP.sub.a and DP.sub.c at which the ascent or descent is starting. Note 
that if Eqs. 3-4 are substituted into Eq. 2, the resultant is that C.sub.f 
will equal zero. Therefore, computing M at the particular ascent/descent 
starting point, results in the correction factor C.sub.f being initialized 
to zero. 
The logic 32 performs a subroutine 98 next to compute the ratio (DP.sub.c 
/DP.sub.a), followed by subroutine 100 which computes the set point value 
for the ratio, P.sub.r as defined in Eq. 3. Note, the value for M computed 
in subroutine 96 is used in Eq. 3. Subroutine 102 is performed next to 
calculate d(P.sub.cd)/dt defined in Eq. 5. A step 104 is then performed to 
compute an error signal indicative of the difference between P.sub.ld and 
P.sub.c during a descent, and P.sub.cc and P.sub.c during an ascent. The 
rate of change of the error signal is then rate limited in a subroutine 
106. The set point for the rate limit of subroutine 106 is the value of 
d(P.sub.cd)/dt computed in subroutine 102. A rate limited error signal is 
then provided to a subroutine 108 which computes an outflow valve command 
to control the variable flow area of the outflow valve 13 (FIG. 1). 
The next step in FIG. 2 is to determine if the descent (or ascent) is 
complete. This determination is performed by a test 110 using the signal 
on the line 31 (FIG. 1) which is indicative of whether the ascent or 
descent has been completed. This signal is typically provided by a 
weight-on-wheels detector or the aircraft flight management system (FMS). 
If the ascent/descent has been completed a return step 112 is executed. 
Otherwise, a subroutine 114 is performed which reads the most recent 
values of P.sub.a, P.sub.c, P.sub.cc, P.sub.cr and P.sub.ld, followed by 
recursive execution at subroutine 98 again. 
An alternative embodiment 116 of the present invention for an aircraft 
descent is illustrated in FIG. 6. The alternative embodiment includes an 
electronic controller 118 which provides the outflow command signal on the 
line 22 to the outflow valve 13. The controller 118 includes a limit 
computation function 120 which receives P.sub.a, P.sub.c, P.sub.cr, 
P.sub.cc and P.sub.ld and provides the adjustable rate limit set point on 
a line 121 to a rate limit function 122. The limit computation function 
120 computes the adjustable rate limit set point using the same equations 
as disclosed in Eqs. 1-5, and can be embodied in either analog or digital 
electronics. A summing function 123 computes the difference between 
P.sub.ld and P.sub.c, and provides a signal indicative thereof on a line 
124. The signal on the line 124 is similar to the error signal computed in 
step 104 (FIG. 2) of the first embodiment. The rate limit function 122 
operates in a well known manner to limit the rate of change of the signal 
on the line 124 if the rate exceeds the magnitude of the adjustable limit 
signal on the line 121. The rate limit function provides a rate limited 
signal on a line 126 to a driver logic function 128 which computes and 
provides the outflow command signal on the line 22 to the outflow valve 
13. 
It should be understood that the scope of this invention is not limited to 
the specific equations presented hereinbefore or to the exact order of the 
steps executed in the flow chart of FIG. 2. Rather the equations are 
readily computable based upon the typical ascent and descent profile of a 
specific aircraft type. Similarly, the steps of FIG. 2 may be performed 
several ways and additions may be made thereto in the practice of the 
present invention. It should also be noted that in the interest of clarity 
in FIG. 3, an assumption was made that the pressure at the takeoff and 
landing points were equal. Clearly the present invention is not so 
limited. 
The present invention will also control cabin pressure at a reduced, linear 
rate even in a situation where the aircraft cabin pressure has to be 
reduced in preparation for landing. Such a situation occurs during 
aircraft approaches to high altitude airports such as Kennedy 
International Airport in LaPaz, Mexico whose elevation is 13,392 feet 
above sea level. In this situation the aircraft cabin altitude of 8,000 
feet during a cruise has to be increased during descent since the landing 
field altitude is 13,392 feet. FIG. 7 is an illustration 130 such a 
descent; the aircraft flies along a curve 132 and the aircraft cabin flies 
along a curve 134. Note, that during the aircraft descent, the cabin 
pressure beginning at a point 136 starts to decrease rather than increase 
in preparation for landing at an airport whose elevation is greater then 
8,000 feet. Similarly during an ascent from a high airport (e.g., LaPaz), 
rather than the cabin pressure being reduced during the climb, cabin 
pressure will be increased since maximum cabin cruising altitude is 8,000 
feet while the departure airport altitude is greater then 8,000 feet. 
The present invention can also be utilized in a closed loop control system, 
rather than the open loop systems of the first and second embodiments 
presented hereinbefore. As illustrated in FIG. 8 a closed loop embodiment 
140 includes a set point computation function 142 which schedules a set 
point for the cabin pressure rate of change and provides a signal 
indicative thereof on a line 144 to an integrator 146. The integrator 
integrates over time to provide a desired cabin pressure signal P.sub.cd 
on a line 147 which is compared to P.sub.c by a summing function 148, and 
an error signal indicative of the difference is provided on a line 150 to 
driver logic 152. The driver logic contains the necessary compensation 
(e.g., proportional and integral paths) to condition the error signal and 
provide a command signal responsive thereto on the line 22. 
Yet another embodiment 154 of the invention is illustrated in FIG. 9. In 
this embodiment the cabin pressure signal P.sub.c on the line 30 is input 
to a derivative/lag filter which provides a signal on a line indicative of 
the rate of change of actual cabin pressure, i.e. d(P.sub.c)/dt. A summing 
function 160 then computes the difference between the actual and desired 
rates of cabin pressure change, and provides an error signal indicative of 
the difference on a line 162. The error signal is input to a driver logic 
function 164 which provides a command signal on the line 22 to the outflow 
valve 13. 
All the foregoing changes and variations are irrelevant to the invention, 
it suffices an adjustable set point for cabin air pressure during ascent 
and descent is computed and used to control the actual aircraft cabin 
pressure rate of change. The adjustable set point value is computed 
throughout the aircraft ascent/descent as a function of P.sub.a, P.sub.ld, 
P.sub.cr, P.sub.cc, P.sub.c and the ascent/descent starting point to 
provide a reduced, linear, cabin pressure rate of change during nonlinear 
aircraft ascents and descents. 
Although the present invention has been shown and described with respect to 
a best mode embodiment thereof, it should be understood by those skilled 
in the art that various other changes, omissions and additions to the form 
and detail thereof, may be made therein without departing from the spirit 
and scope of the invention.