GB2046690A

GB2046690A - Improvements in or Relating to Instrumentation Apparatus

Info

Publication number: GB2046690A
Application number: GB8002960A
Authority: GB
Original assignee: UK Secretary of State for Defence
Current assignee: UK Secretary of State for Defence
Priority date: 1979-02-01
Filing date: 1980-01-29
Publication date: 1980-11-19
Also published as: GB2046690B

Abstract

This invention relates to instrumentation apparatus for aircraft and provides an indication of increased aerofoil drag for rotary-wing aircraft in icing or potential icing conditions. The apparatus is arranged to receive a set of parameters relating to external conditions and a set of parameters relating to internal conditions and includes means for computing therefrom a value for rotor torque demand for the conditions defined and for determining the difference between this demand and the torque actually transmitted to the rotor. The apparatus may be arranged to provide an indication of any increment of rotor torque required to fly the aircraft over the computed rotor torque to thereby indicate and warn of the possibility of rotor icing, and to provide a signal to initiate de- icing. External parameters used are air temperature and pressure. Internal parameters are aircraft speed, rotor collective pitch angle.

Description

SPECIFICATION Improvements in or Relating to Instrumentation Apparatus This invention relates to instrumentation apparatus for aircraft.

The problem of ice accretion on aircraft aerofoil surfaces is one of the most acute yet to be solved in the field of aviation.

Particularly affected by icing are rotary-wing aircraft where the consequences of severe icebuild-up are potentially disastrous. Problems are caused because ice forms on the aircraft structure, which can endanger engines, block ducts, obscure windscreens, and lead to an increase in parasite drag. However, the most critical effect can be the large and rapid increase in drag coefficient of the main rotor aerofoils caused by icing of their leading edges. This may produce power demands from the engines that exceed the capability of the transmission system, or affect the rotor blade pitching moment and overstress the control system. The drag effect can also be such as to cause loss of autorotative capability.

Some instrumentation is available to detect icing in airframe structure but icing on the rotor cannot at present be detected readily other than by indirect methods such as stress monitoring.

Increased drag coefficient means that a greater torque must be supplied to the rotor than would be required in clear air (ice-free) conditions. In order to try and identify this condition in an aircraft flying in potentially icing conditions a crew member may monitor delivered torque readings and compare them with tables giving values obtained in clear air flight. However this proves to be laborious and unreliable because of dependence of aircraft weight, external configuration, and airspeed, plus the need to be in steady level flight which can be difficult to attain in icing cloud.

The practical solution to this problem is the strict imposition of much restricted operational limitations on aircraft if a chance icing exists.

This solution is not entirely effective because weather conditions are difficult to predict with sufficient precision, and particularly in an emergency, a decision may be taken to fly in marginal conditions with potentially disastrous results. This situation may also arise as a result of a sudden change in the weather, and the character of icing clouds particularly has been wound to change dramatically over short distances.

It has been discovered that a value for the torque required for a particular flight configuration in clear air may be predicted from collective pitch (that is rotor aerofoil angle of attack) and other parameters. It is an object of the present invention to utilise this discovery such that aviation safety may be improved.

According to the present invention instrumentation apparatus for use with aircraft having a driven rotor includes means for receiving a set of parameters relating to external conditions, means for receiving a set of parameters relating to internal conditions, including collective pitch, means for computing from these parameters and in accordance with a predetermined law a value for a rotor torque demand for the conditions defined, means for receiving a set of parameters relating to torque transmitted to the rotor and means for determining the difference between the computed torque demand and the torque transmitted.

It will be realised that the invention provides an indication of the increment of torque required to fly an aircraft in a particular configuration over the torque that should be demanded by that configuration in the conditions defined by the received parameters assuming no increase in drag coefficient, and hence provides a warning of increased aerofoil drag.

A display of torque increment, in accordance with the present invention may advantageously be provided as a pilot monitored instrument. This enables the pilot of an aircraft to appreciate the existence of rotor icing, to assess its rate of increase and to initiate appropriate action. A particular advantage of the present invention is that it allows continuous monitoring in potential icing situations in all flight configurations and reliability is only slightly affected by gentle manoeuvres (including climbs and descents) or changes in aircraft weight (such as when fuel is consumed).

In helicopters where a recording system for performance measurement and other reasons, exists, information derived from a plurality of transducers is fed to a central recording unit. This information is transferred in serial binary format via a time division multiplex information link. In accordance with the present invention instrumentation apparatus may be advantageously interfaced with such a link to receive a set of parameters relating to external conditions and a set of parameters relating to the actual torque transmitted to the rotor. Where no such recording unit exists, the provision of suitable transducers is straightforward.

Preferably information is received at an input port to a micro-computer including a Central Processor Unit and random access memory (RAM) as means for computing a value for clean rotor torque for the conditions. Such computation may be performed according to the instruction set of the Central Processor Unit by transfer of information to and from RAM. Information relating to the actual torque transmitted to the rotor may also be received at an input port to the micro-computer and additional computation performed to provide an indication of the difference between computed clean rotor torque and actual rotor torque at an output port. This output may advantageously be converted from binary to analogue format to provide a single electrical output for transmission to recording apparatus or a meter giving a real time indication of rotor aerofoil condition.Alternatively the output may be used as an input signal to a de-icing or other control system.

In order that features of the invention and its operation may be better appreciated the following example is presented.

During investigations of helicopters in icing conditions, the following set of parameters have been identified as useful for computation of a value for clean rotor torque: CpCollective pitch angle, degrees H-Prnssure altitude, feet T-External static air temperature, degree C V-Indicated Airspeed, knots.

The computation of a value for clean rotor torque (0c) may advantageously be performed in stages by deriving the following parameters from the above.

6--relative pressure a-relative density where: in accordance with well known principles:

=1.0-0.36 ( Hp +0.00048 ( Hp ) 2 1000 ) 1000 clllu.

2888 a (273+T) whence: Q=-20+( 6-0.0008V2) (Cp'y'a-3)+0.363 (Cpa-3) Indication of torque increment (AQ) proceeds as follows: For a twin-engined helicopter the actual torque transmitted to the rotor (QT) is given by: Q.+Q2 Qffl 2 where Q1 is No 1 engine measured torque expressed as a percentage Q2 is No 2 engine measured torque, hence: AQ QT QC It will be appreciated that the above example shows how instrumentation apparatus in accordance with the present invention may be configured or programmed to operate by, for example, the means hereinbefore described.

Claims

Claims The matter for which the applicant seeks protection is:

1. Instrumentation apparatus for use with aircraft having a driven rotor including means for receiving a set of parameters relating to external conditions, means for receiving a set of parameters relating to internal conditions, including collective pitch, means for computing from these parameters and in accordance with a predetermined law a value for rotor torque demand for the conditions defined, means for receiving a set of parameters relating to torque transmitted to the rotor and means for determining the difference between the computed torque demand and the torque transmitted.

2. Instrumentation apparatus as claimed in claim 1 and arranged to provide an indication of any increment of rotor torque required to fly the aircraft over the computed rotor torque.

3. Instrumentation apparatus as claimed in claim 1 or claim 2 and interfaced with an information link between a plurality of transducers and the aircraft central recording unit.

4. Instrumentation apparatus as claimed in claim 1, claim 2 or claim 3 and wherein parameters are received at an input point to a micro-computer.

5. Instrumentation apparatus as claimed in any one preceding claim wherein the predetermined law is or an approximation to: Q,=--20+(6-0.0008V2) (Cp""a-3)+0.363(CpV-3) where Qc, V, Cp, and a are as hereinbefore defined.

6. An aircraft having a driven rotor and including instrumentation apparatus as claimed in any one preceding claim.

7. A helicopter including instrumentation apparatus substantially as hereinabove described.