Feedback laws based upon optimal control theory were derived.
and these resulted in a reduction of the structural loads
on the wing of a simulated aircraft. Various models of the
aircraft dynamics were used. the most complete being of order
79. This model included rigid body motion, structural flexibility
effects, unsteady aerodynamics, gust dynamics and actuator
dynamics. The structural effects were characterised by
the first fifteen bending modes. The subject aircraft studied,
was considered to employ active ailerons and elevators and was
subjected to manoeuvre commands and simulated atmospheric
Extensive numerical tests have shown that feedback laws
derived from reduced dimension models performed comparably with
the feedback law based on the most complete model. Tests were
made on feedback laws ranging from order 79 to order 5. It was.
however. not possible to reduce the number of feedback variables
below five as this then affected the stability of the aircraft.
The law based upon five state variable feedback was given the
designation 'safety law'.
One of the consequences of operating under the action of
the 'safety law' was that the same level of load reduction
could not be achieved as was obtained whenever a full state
feedback law was employed. In addition 'safety law' operation
was often marked by large transient oscillations of the wing
root bending moment and it was considered that this would
subsequently affect the fatigue life of the structure. An
observer design was then investigated which reconstructed the
complete state vector from a selection of measurements of the
sensor signals appropriate to the 'safety law'. Results have
shown that it is possible to achieve a practical implementation
of such a scheme which will possess all the attendant advantages
of full state feedback control.
A consequence of reducing the strength of the wing of
the aircraft as a result of employing an active load alleviation
scheme is that a considerable degree of reliability of
the control system, higher than that of both the basic airframe
and its propulsive system, will be required. Because the use
of hardware redundancy techniques as a means of providing the
required degree of reliability would be expensive, software
redundancy techniques suggest an attractive alternative. One
example of how software redundancy may be employed is demonstrated
in respect of , checking the analogue feedback gain
controller used in the aircraft to implement linear feedback.
It is shown how a-microprocessor may be effectively employed
to introduce a surrogate gain should one or more of the channels
of the controller fail.
A Doctoral Thesis. Submitted in partial fulfilment of the requirements for the award of Doctor of Philosophy of Loughborough University.