Velocity vector control system augmented with direct lift control

A pilot controlled stability control system that employs direct lift control (spoiler control) with elevator control to control the flight path angle of an aircraft. A computer on the aircraft generates an elevator control signal and a spoiler control signal, using a pilot-controlled pitch control signal and pitch rate, vertical velocity, roll angle, groundspeed, engine pressure ratio and vertical acceleration signals which are generated on the aircraft. The direct lift control by the aircraft spoilers improves the response of the aircraft flight path angle and provides short term flight path stabilization against environmental disturbances.

BACKGROUND OF THE INVENTION 
The invention relates generally to aircraft control systems and more 
particularly concerns an aircraft flight path angle control system that 
employs direct lift control. 
The flight path angle is the angle between the aircraft's inertial velocity 
vector and the local horizontal reference plane. In the past, 
pilot-controlled Stability Augmentation Systems (SAS) designed to control 
the flight path angle have utilized the aircraft's elevators (or 
horizontal stabilizers) for the control. The main disadvantage with these 
systems is there is a delay between the time the pilot changes the 
position of the elevators and the time the flight path angle changes in 
response thereto. Also with elevator control there is limited ability to 
counteract short term flight path fluctuation due to environmental 
disturbances. 
It is the primary object of this invention to provide a pilot-controlled 
SAS that decreases the delay between the time that the pilot actuates his 
control until the time that the flight path angle changes in response 
thereto. 
Another object of this invention is to provide short term flight path angle 
stabilization due to environmental disturbances. 
Other objects and advantages of this invention will become apparent 
hereinafter and in the drawings. 
SUMMARY OF THE INVENTION 
This invention is a pilot controlled SAS that employs direct lift control 
in conjunction with the aircraft elevators to control the flight path 
angle. The direct lift control provides lift quickening and thus improves 
the response of the aircraft flight path to pilot commands. Also the 
direct lift control augments elevator control in providing short term 
flight path stabilization due to environmental disturbances. Long term 
flight path control is maintained by a high performance closed loop 
elevator control system. 
A transducer generates a pitch control signal in response to pilot inputs 
and instruments on the aircraft generate signals proportional to pitch 
rate, vertical velocity, groundspeed, and vertical acceleration. A 
computer on board the aircraft receives these signals and generates an 
elevator control signal and a direct lift control signal for controlling 
the elevators and spoilers on the aircraft. In addition, a roll angle 
signal is generated on the aircraft and applied to the computer to cancel 
a loss of the vertical component of lift due to bank angles.

DETAILED DESCRIPTION OF THE INVENTION 
Turning now to the embodiment of the invention selected for illustration in 
the drawings the number 11 in FIG. 1 designates a pilot controlled pitch 
transducer which generates a pitch control signal X. In operating this 
invention the pilot views the flight path angle on the aircraft cockpit 
display and he changes the value of the pitch control signal X to change 
the flight path angle to some new commanded value. This invention quickens 
the response time between the time of a change of the pitch control signal 
X and the time the change is shown on the aircraft display as a change in 
the flight path angle. Moreover, this invention provides a well-damped, 
highly stable response. 
Instrumentation on the aircraft generate feedback control signals: a pitch 
rate gyro 12 generates a pitch rate signal .theta.; an air data computer 
13 generates a vertical velocity signal h; an aircraft engine 
instrumentation 15 generates an EPR signal; and an inertial navigation 
system 14 generates a groundspeed signal V.sub.g, a vertical acceleration 
signal h, and a roll angle signal .phi.. The pitch control, pitch rate, 
vertical velocity, groundspeed, vertical acceleration, roll angle and EPR 
signals are all applied to analog-to-digital converters 16 where they are 
converted to digital signals. These digital signals are applied to a 
flight control computer 17 which generates an elevator control signal and 
a spoiler control signal. These elevator control and spoiler control 
digital signals are applied to digital-to-analog converters 18 which 
convert them to analog signals. The analog elevator control signal is 
applied to elevator power control units 19 to control the elevators 20 on 
the aircraft; and the analog spoiler control signal is applied to spoiler 
power control units 21 to control the spoilers 22. 
Flight control computer 17 as shown in FIG. 2 includes a divider 30 which 
receives the vertical velocity and groundspeed signals and generates a 
flight path angle signal .gamma. by dividing the groundspeed signal into 
the vertical velocity signal. The pitch control signal X from the 
pilot-controlled transducer 11 is applied through a gain and noise filter 
31 to a summing device 32. The output of summing device 32 is the elevator 
control signal. The signal from the gain and noise filter 31 initiates 
pitch response through the elevators. The characteristics of the gain and 
noise filter 31 is defined by the expression: 
##EQU1## 
where K.sub.PDE is a constant determined by the characteristics of the 
aircraft on which the system is used. .tau..sub.2 is a time constant and s 
is a Laplace operator. The pitch control signal X is also applied to a 
constant multiplier 33 which multiplies X with a constant K.sub.PSP. The 
output multiplier 33 is applied through a summing device 34 to a spoiler 
command limiter 35 the output of which is the spoiler control signal. The 
signal at the output of limiter 35 provides an immediate lift increment to 
start changing the flight path angle in a favorable direction. To 
counteract the spoiler pitching moment, the spoiler control signal at the 
output of limiter 35 is cross fed through a constant multiplier 36 which 
multiplies the signal by a constant K.sub.CF to a summing device 32 to 
provide additional elevator control. 
At the instant the system is activated (turned on) by the pilot the flight 
path angle signal .gamma. at the output of divider 31 is applied to an 
integrator 37 to set the integrator to the initial value of .gamma.. The 
pitch control signal X is then integrated by the integrator 37 to form the 
commanded flight path angle signal .gamma..sub.c at the output of the 
integrator. This signal is compared with the computed .lambda. signal at 
the output of divider 30 by means of a summing device 38 to produce an 
error signal .DELTA..sub.65. 
The primary elevator stabilization signal is derived by passing the error 
signal .DELTA..sub..gamma. through a high gain lead lag filter 39 the 
output of which is applied to summing device 32. The characteristic of 
filter 39 is defined by the expression: 
##EQU2## 
where K.sub.1, K.sub.2 and K.sub.3 are constants. This part of the 
elevator control signal is the primary elevator stabilization signal. 
Filter 39 is the part of the invention that optimizes stability and 
response. The .DELTA..sub..lambda. signal is also applied through an 
integrator 40 the output of which is applied to the summing device 32. 
This part of the elevator control signal takes care of possible 
steady-state standoff errors which would occur due to bias error signals 
or elevator trim requirements. In addition the signal .DELTA..sub..gamma. 
is applied through a constant multiplier 41, which multiplies the signal 
by a constant K.sub..DELTA..gamma., to summing device 34 to provide long 
term spoiler corrections for the flight path angle errors. The vertical 
acceleration signal h from the inertial navigation system 14 is applied 
through a constant multiplier 42, which multiplies the signal by a 
constant K.sub.h, to the summing device 34. This part of the spoiler 
control signal essentially provides a .gamma. signal for stabilization. 
The pitch rate signal .theta. from the pitch rate gyro 12 is applied 
through a constant multiplier 49, which multiplies the signal by a 
constant K.sub..theta., to a washout filter 43. The characteristic of 
filter 43 is defined by the expression: 
##EQU3## 
where .tau..sub.1 is a time constant. The output of filter 43 is applied 
to summing device 32 to provide short period mode dampening in the 
elevator control signal. The roll angle signal .phi. from the inertial 
navigation system 14 is squared by a multiplier 44 and applied through a 
constant multiplier 45 to the summing device 32. Multiplier 45 multiplies 
the roll angle signal by a constant K.sub..phi. to cancel a loss of the 
vertical component of lift due to bank angles. 
Another feature of the invention which is claimed in a co-pending 
application is the use of the EPR feedback for cancelling pitching moments 
due to thrust changes. The technique is based upon knowledge of the 
relationship between engine location, engine thrust, EPR, and elevator 
effectiveness. An analysis of these factors produces a gain of K.sub.EPR 
which when applied to the EPR feedback signal commands the proper amount 
of elevator to cancel thrust induced pitching moments. To implement this 
part of the elevator control signal the EPR signal initially sets the 
reference EPR 46 at the time the pilot engages the system. Thereafter the 
generated EPR signal is compared with the reference EPR signal by means of 
a summing device 47. The difference output of summing device 47 is 
multiplied by a constant K.sub.EPR by means of a constant multiplier 48 
and then applied to summing device 32. Two benefits are immediately 
available from this scheme: pitch disturbances due to thrust changes are 
cancelled, and an elevator bias signal is provided downstream of the 
washout integrator 40 allowing a reduction in the integrator gain and 
thereby contributing to an increase in system stability. 
All of the constants shown in the block diagram in FIG. 2 are determined by 
the particular aircraft on which the invention is used. A good estimate of 
each constant can be determined from the available data on the aircraft 
and thereafter the constants can be adjusted to obtain the desired 
responses. 
This invention has been used on a flight simulator by Langley Research 
Center in Hampton, Va. The simulator represents a twin-engine medium jet 
transport that was modified to include an advanced research cockpit, 
direct lift control capability, and onboard flight research equipment. The 
values of the constants in FIG. 2 that were used are as follows: 
K.sub.h =4.9 
K.sub.PSP =2.4 
Limiter 35=.+-.8.degree. 
K.sub.cf =0.35 
K.sub.c =0.33 
K.sub..DELTA..gamma. =8.0 
K.sub.PDE =1.3 
.tau..sub.2 =0.09 
K.sub.1 =20.0 
K.sub.2 =0.8 
K.sub.3 =2.5 
K.sub.I =0.3 
K.sub..theta. =4.0 
.tau..sub.1 =16 
K.sub..phi. =0.004 
K.sub.EPR =8.2 
The advantages of this invention over previous pilot controlled stability 
augmentation systems are numerous. It provides lift quickening and thus 
improves the response of the aircraft flight path angle to pilot commands; 
it provides short term flight path stabilization against environmental 
disturbances; it improves the ride quality of the aircraft; it improves 
the pilot accuracy in tracking a glide slope; and it improves flare and 
touchdown performance of the aircraft.