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1. k a 4 Yy 1 N 4 siine ERRE UTS REPT LTT IT ETE PTE TE TST TIT TT TH EE TT T Ce se ae i n measured and corredated to aircraft behavior as recorded by the test instrumentation system As was the case with actuator signal measurement stability system electrical signals should be carefully assessed for requirement of electrical isolation of the measurement 7 The stability system output is usually an electrical signal to an actuator placed in the flight control system The voltage can be measured directly and through calibration or specification data the actuator motion can be determined A more common practice is to measure the actuator motion with a position transducer 7 PONS SYSTEMS Current helicopter practice is to use weapon systems that are an integral part of the design as opposed to earlier vehicles where the systems were added in the field as required This development allows installation of more powerful systems and at the same time gives an opportunity to minimize the weapons effects on the flight vehicle However greater complexity usually accompanies the increased capability and in turn more instru mentation is needed to measure and record the data The total weapon system includes sighting and aiming controls and the weapon The weapon system is usually developed and tested independently of the aircraft This effort is concerned with assuring the system will deliver the specified ordnance wi
2. Engine Vibration Varies by Piezoelectric location ac elerometer and axis 4 1 1 Pitch Attitude 4 1 1 Roll Attitude 4 1 2 Yaw Attitude Angular Pitch deg sec 0 1 Rate EN 180 max but can be reduced for in creased resolution EOE ee tat Angular Roll deg sec 100 Rate o n Angular Yaw deg sec t60 1 Rate Linear Z 4 1 t0 1 0 001 Acceleration X and Y t2 4 5 Airframe Vibration G 0 001 Can include both low and high freq u O to 100 t5 0 5 t2 uency accelerometers 100 should be 10 deg 0 to 100 O to 100 t2 lt Structural Loads above expected yield point 100 should be 10 above the maximum calculated a D _ Q amp Blade Positions Flight Control Positions Flight Control lb Forces Actuator Dosition Swasii Plate Angle Long Lat Range dependent on aircraft and seci fication require amp Dir 10 Pedal 25 0 to 100 in Bs SOE RR tee ci pee ene oo deg to 100 100 is maximum flight Control input Se tana a btn aD RA Sati tha ca rls i SIA EP Perey FETS IS TG 1 f e a EE PAEL u we re ci aS Teo Perf ES MRE ET aR ap ete SES A2 1 APPENDIX II TYPICAL INSTRUCTIONS FOR DEVELOPING AND MAINTAINING RECORDED INSTRUMENT PARAMETER LIST This appendix p
3. 4 7 3 Internal Noise The internal noise level in the helicopter musc be measured to evaluate crew contort performance communication and safety aspects Consideration must be given to measurement at point locations such as the pilot s ear or to obtaining data needed to create noise profiles Noise data for the passenger section are of particular interest sicee these personnel do not normally wear helmets or protective gear Selection of the sensors and recording equipment must accommodate a frequency range from 20 to 10 000 HZ and overall decibel levels up to 120 This will be influenced by windows rotor and engine speed variations and any weapons firing Sensor and recording requirements will be further addressed in the far field measurements Section 8 2 ES OE APS b ROTORS AND PROPELLERS Rotors and conventional fixed wing aircraft propellers have a great deal in common however rotors have several features which render them considerably more difficult s to instrument and test Significant differences are xI a Rotors have blades which are longer and thinner with less rigidity b Rotors are controlled with both collective and cyclic pitch inputs c Blades may be attached to the hub with various hinge arrangements d Rotors have different and complex axial and in plane flow relations j ELE SOR AE CNS LIES The critical nature of the rotor system dictates that a great deal of data and 4 analysis be consider
4. ture sensitive material A change in color shows that a specified temperature has been exceeded This material is available as a paint or as a template A wide range and choice of increments can be selected to meet the expected requirement More accurate data can he obtained from thermocouples bonded to the test surface Compartment temperature is measured wivh a resistance thermometer using the techniques discussed for free air temperature measurements 3 6 Engine Pressure Engine power and compressor performance information requires measurement oy compressor discharge pressure Engine specifications will provide the compressor ratio data which is used to select a sensor with the proper range Exhaust gas pressure is measured with individual sensors or with an array of sensors on a rake as discussed for the engine inlet Care should be taken to insure that the sensor chosen is operationally compatible with the installation as well as suitable for the gas temperature being measured 3 7 Fuel Fuel measurements include flow rate temperature and quantity used Flow rate and temperature are necessary for engine or aircraft performance tests while quantity used is needed to determine aircraft weight A careful study of the fuel system should be made prior to designing the flow measurement instrumentation Special attention must be given to relative flow from various tanks fuel transfer for CG control fuel by pass valves or vapor return lines Fuel
5. ATMOSPHERIC 2 1 Air Data Instrumentation All flight tests require some meastremenrs of atmospheric data Measurements inviude pressure teuperature liquid water cortent dust or cetris concentrations humidity and flow angles The measurements may be devoted to the far field relative to the aircraft or local conditions at a component or surface Special problems arise during climb and descent or dynamic maneuvers nesr the ground in various surface winds In the latter cage a ground station is often used to define the far field environment 2 1 1 Free Air Temperature Frcs air temperature systems must be installed so that they will receive a minimum influence from the aircraft The sensor should be shielded from heat generating sources or trom hot airflow Solar radiation should also be considered Common practive is vo install a calibrated test system to record the data This test system is then used as the standard to evaluate the basic aircraft temperature sensing installation The test sensor is usually mounted on the airspeed boom When a boom is not available the sensor is ofteu mounted on the underside of the nose of the aircraft Many test sensors have a de ice capability however care must be used to insure that the de ice is on only at the specified conditions Typically operation above 0 C or below 30 m s 59 Kn will introduce a 2 C error The activation of the de ice may be manual or auto matic The system must incluce a co
6. and lifting surfaces The input may be modified or shaped during the process Aerodynamics of the rotor are continually providing feedback throughout the system Aircraft motions provide pilot cues and those motions may also cause stability systems to respond The instrumentation may simply measure the pilot actions or it may be required to measure each input and motion of every system component When instrumenting the control system special care must be used to insure that the instruments do not introduce forces cause inter ference or change the characteristics even if the test system should fail 6 1 Cockpit Controls Basic helicopter cockpit controls are cyclic stick for longitudinal and lateral control pedals for directional control and a collective lever for thrust control When free those controls are essentially cantilever beams and will vibrate as driven by the airframe or control system feedback forces Linear accelerometers are placed at the top of the cyclic control and oriented to measure longitudinal and lateral vibrations An accelero meter is placed on the collective to measure vertical vibrations Similar accelerometers are placed on the pedals to measure longitudinal accelerations Cockpit control forces are usually not measured directly unless a specially instrumented hand grip or hand held force gage is used Normal practice is to mount strain gages on the control rod attached to the end opposite the hand grip or foot pedal Th
7. 2 Engine Torque 15 3 3 Shaft Torque 15 3 4 Inlet 15 3 4 1 Inlet Precsure 16 3 4 2 Inlet Temperature 17 3 4 3 Inlet Devices 17 3 5 Engine Temperature 17 3 6 Engine Pressure 18 3 7 Fuel 18 3 7 1 Fuel Flow 18 3 7 2 Fuel Temperature 18 3 7 3 Fuel Quantity 19 3 8 Power Extraction 19 3 9 Power Plant Controls 19 3 9 1 Cockpit Controla 19 3 9 2 Engine Controls 19 3 10 Engine Vibrations 20 4 AIRFRAME 20 4 1 Attitude 20 4 1 1 Pitch and Roll Attitude 20 4 1 2 Yaw Attitude 21 4 2 Angular Rate 21 4 3 Angular Acceleration 21 4 4 Linear Acceleration 21 4 5 Vibration 21 4 5 1 Sensor Location 22 4 5 2 Sensors 22 4 6 Loads 22 4 6 1 Sensor Location 22 4 6 3 Sensors 22 4 7 Cockpit and Cabin Environment 23 4 7 1 Air Temperature and Airflow 23 4 7 2 Surface Temperature 23 4 7 3 Internal Noise 23 5 ROTORS AND PROPELLOKS 23 5 1 Blades a4 8 1 1 Airflow 24 5 1 2 Blade Positions 24 6 2 Hubs 24 5 3 Pitch Links 24 8 4 Data Transfer 25 8 5 Non Rotating surfaces 26 Document provided by SpaceAge Contr Inc http spaceagecontrol com nace tl PTAS A E es ERLE erates E EE a Lae ae ee eT ee a esate SSS YES a pa a a Aaeeea oa emrenin e toc QOGUMENt provided by SpaceAge Control inc 6 FLIGHT CONTROL SYSTEM 6 1 Cockpit Controls 6 2 Mechanical Linkages 6 3 Actuators 6 4 Swash Plate 6 5 Stability Augmentation Systems 7 WEAPONS S
8. Angles of Attack and Sideslip 2 2 2 Relative Wind deg 180 1 0 Vertical Relative Wind deg 180 1 0 Azimuth Omni directional 50 in all wongitudinal may be Airspeed een er extended to 250 3 1 1 Gas Generator 50 to 110 Frequency or period counting digital tach with 10 sec quartz clock refe rence Speed Shaft Speed 0 to 100 Transducer used is often the normal aijroraft torque Shaft Torque J0 to syatem 3 4 1 Inlet Pressure 1b in ti psid Accuracy includes error of pressure rake ESE EES pee ween ae ces Swart ses SET tee Pe oe iD SRS ob hea aR ae RiP et ee eee Bie Sten REFER RESO ENCE PARAMETER ACCURACY LUTION REMARKS Inlet 230 C Temperature i Engine 900 C 3 0 Over normal operat Temperature ing range Engine Pressure 1b in 150 psig Bleed Air used primarily with Fuel Flow gal hr Variable by acft type turbine sensor Fuel Used Variable by acft A type Electromotive Volts O to 100 1 0 Force Electric Current peres 0 to 100 deg O to 100 Volume measurement Volume measurement used primarily with turbine sensor A C power source generally 28VNC or 115Vac ee Stan ates ened Cea BEE ore ni fee ot A A C power source generally 28VDC or 115VAC Generally repeat able to ti Cockpit Power Controls wa To Mat ee em ale
9. Most engines have an integral rotational speed sensor which provides an electrical signal whose frequency is propor tional to the speed This signal is used to drive standard cockpit instruments and can algo be input to a test recording system The standard instrument is usually not suitable for recording test data from the cockpit and is often replaced or paralleled with a high resolution test instrument This method is most useful for the pilot when conducting tests or i use by observers monitoring test progress for completeness or quality of test results Power turbine speed measurement is often tachometer generator system similar to that described for the compressor speed and comparable methods are used The power turbine shaft may be directly available or may have an integral gear reduction ransmission The most convenient shaft is fitted with speed measurement instruments and if necessary gear Document provided by SpaceAge Control Inc http spaceagecontrol com p aeta PESEN AA IRTIR T VAVA SAA MAA REPETE AIEE ene MAREE PEE Aa P Dindaia et PEAN P r P PPE PET EET TE E EE E paea SONS i gt k JaA Fen Sean TEATS EST TS ren a PES be NTI EI OT AT OO Nt ATE a TRG teea Fana arcas nt taila ind SETS CLS KL riaan Sira oct he PRERE EN R SERRA W W STA OET BENE REF AR ratios are used to calculate the turbine speed A typical test instrument to measu
10. a matrix of small holes drilled in the blade and tubing is then used to duct the static pressure to differential pressure transducers Physical alteration of the blade must be accomplished under the direction of structural engineers The tubing should be as short as possible to reduce lag in the system and to provide the best response to rapid pressure changes In order to determine the pressure distribution accurately such a large number of pressures must be sensed that the recording capability is often overloaded Commutation is used to reduce the number of recording channels The speed of commutation and the type of data must be such that interruptions do not invalidate the results The static pressure changes are usually small values and high sensitivity is needed i 5 1 2 Blade Positions The blade positions aro controlled by inputs of collective pitch to all blades gnd cyclic pitch which varies as the blade azimuth changes Aerodynamic forces cause vertical blade motions flapping in plane motions lag and torsional pitching motions The sum of those motions combine with rotational speed and free stream air to produce a local blade angle of attack The blade ungle of attack is different at each blade section and is changing very rapidly Blade angle of attack is not measured directly but can be calculated from the pressure distribution data For articulated or teetering rotors blade flapping is measured at the blade hub attachment with a
11. amplitudes and frequencies can vary widely Basic frequencies will be multiples of the main rotor speed Superimposed will be the tail rotor frequencies as well as those from structural components and other rotating parts The main rotor will generate in plane and vertical vibrations Fuselage vibration absorbers may be used These absorbers are usually effective only within a certain frequency range During operation at other frequencies they may amplify the basic vibration Aircraft compartments usually have no environmental control and instrumentation placed there will experience a variety of conditions Where there is no heating the compartment temperature wiil vary from 40 C in desert conditions to 25 C during high altitude tests When instruments are placed in compartments near engines or transmissions special care must be used to determine compartment temperatures prior to installation A marine environment leads to consideration of any salt spray that may occur Tests in a desert situation generate dust and debris from the rotor wash This dust can be very fine and dense and will probably enter any compartments not specialiy sealed In addition to the cold temperatures ice on the rotor and airframe can signifi cantly change the vibratiou environment of sensors and recorders Weapons firing tests generate severe local pressure variations and alter the aircraft vibration character istics Special landing tests such as minimum distance over an obst
12. external area may require obtaining samples for laboratory analysis Gas temperature is measured with thermocouples 8 GRO SUPPORT_INS NTATION E5 Helicopter missions involve a relatively large amount of time spent near the ground at hover and low speeds and it is necessary to accomplish much of the testing in a similar environment in hover the helicopter may contaminate the near field atmosphere such that airborne measurementa are inaccurate or unreliable Hover performance is most easily and safely conducted by measuring thrust rather than loading the aircraft with weight During takeoff and landing maneuvers or near ground maneuvers the flight path must be corrected to a zero wind condition which requires measurement of ground distances and wind velocities Independent data recording systems may be used and can cause extremely difficult data correlation problema 8 1 Atmosphere Atmospheric measurements include wind speed and direction ambient air tempera ture preasure and humidity The measurements should be taken as near to the helicopter as possible while ensuring that the air mass is undisturbed by the downwash Temperature and wind conditions usually vary with height above the ground so that a ground surface measurement does not describe the helicopter operating environment Best results are obtained when the sensors are mounted on a tower at various heights above the ground The data then gives a profile of the gradients inve
13. i Y Tee Sa YES SES EURA E BEES GE TIES TS E i oe rennin a e vi its wens crinean A QcuMeEnt pravided by spaceAge Cant HIGHER PRESSURE IS RETR DUE TO THE LOWER WIND AIRSPEED THROUGH v THE RETREATING Vw VENTURI LOW PRESSURE IS INDUCED n IN THE ADVANCING ARM DUE TO THE INCREASED SPEED THROUGH THE ADVANCING VENTURI Figure 2 2 2 4 LORAS Airspeed Sensor e Rosemount The Rosemount orthogonal airspeed sensor is manufactured by Rosemount Engineering Company Minneapolis Minnesota The system includes a sensor airspeed indicator transducer analog multiplier unit and tubing Sensor dimensions and operation is outlined in Figure 2 2 2 5 ELECTRICAL CONNECTOR SENSING PORTS THIS AREA HEATED SECTION am 32 a 18 WATTS iN FORWARD peceeune TS LEFT RIGHT oY Y SECTION A X AFT Figure 2 2 2 5 Rosemount Airspeed Sensor The sensor contains internal electrical wiring for deicing Power consumption with deicing operations is 250 watts in flight and 150 watts in still air The Rosemount orthogonal airapeed indicator is a dual pointer instrument with both pointers moving rectilinearly The dual traversing pointers are driven by DC sigrals from the Rosemount transducer and move the pointers through scales representing 60 Kn 30 m s forward to 40 Kn 20 m s aft and 50 Kn 25 m s left to 50 Kn 25 m s right respe
14. large height errors in the computations and the data is not useable Height above ground is best obtained from a radar altimeter Extreme maneuvers may cause radar altimeter problems in which case a precision pressure altimeter should be considered The system should be field calibrated before each test The RU is placed 1000 m 3300 ft from the aircraft and the range calibrate screws on the DMU are adjusted until the correct range is displayed on the unit The system has an adjustable measuring rate The test aircraft may have a doppler or Inertial Navigation System INS which can be used to obtain accurate space position information Ref 31 32 33 34 and 35 Such a system may be used in conjunction with an instrumented range and since it is self contained within the aircraft may be useful for remote site operation The navigation system may be a stand alone system however it is more common to have a central unit which interfaces with other aircraft systems The system must be carefully analyzed to determine what information is available from the system and how this data can be obtained without altering the system operation The navigation system measures component velocity in the aircraft axes and in conjunction with a heading input ccmputes ground speed and direction Inputs from the airspeod system are then compared with ground speed to obtain wind information Some systems also make navigation corrections for attitudes angles of a
15. often will include ground operation to inuure that natural frequencies of the airframe are different than rotor excitation frequencies These types of test are mainly concerned with structural integrity The flight tests are generally concerned with crew or passenger environment or determination of conditions to be experienced by avionics aircraft systems or cargo 4 5 1 Sensor Locations Sensors should be located in all areas where vibrations can be transmitted to the crew members Typical location are the seats flight controls instrument panels foot rests and consoles in the cockpit Each potential location must be analyzed to determine the number and orientation of the required sensors Consideration should also be given to instrumenting external stores pylons doors horizontal and vertical tail surfaces It may also be necessary to obtain data for wings and landing gear Vibration of aircraft components is measured and compared with excitations to evaluate performance of the mounting or damping mechanisms Mounting of the sensors will vary according to the physical makeup of the accelerometer and the mounting location In all cases however the vibra tional characteristics of the structure under test should be altered as little as possible Care should be taken to insure that the natural frequency of the accelerometer is not shifted into the frequency range of interest by the mounting technique This can happen quite easily if poor surfa
16. on the shaft and the brushes mounted on the stationary airframe This arrangement usually gives poor performance because the main rotor shaft may move independently from the airframe which will affect the brush slip ring performance This can be avoided by mounting the brushes on a stationary standpipe mounted inside the rotor shaft With this installation there is no relative movement between the brushes and the slip rings Wiring is then routed inside the standpipe through the transmission and then to the data recorder A typical standpipe installation is shown in Figure 5 4 1 INSTRUMENTATION SEALED BEARING le a ENGINE lt TRANSMISSION CASE NTT a a a DT T T X ATTACHED BY PIN SCRE TO DATA RECORDING OR SNAP RING SYSTEM Figure 5 4 1 Typical Rotor Shaft Standpipe Installation An alternate method to transfer the data is to use a telemetry system A trans mitter receiving power inductively from stationary coils is mounted on the shaft and a receiver is stationary on the airframe The distance between the two should be minimized to reduce transmitter size and power requirements Extreme care must be used so that shaft balance problems are not introduced by the transmitter installation Some advantages of telemetry compared to slip rings are a System can be designed for variety of installations whereas a slip ring is usually unique to one application b
17. other desired temperature This igs accomplished through the resistance temperature relationship for platinum which is given by the Callendar Van Dusen equation Ry T T T tr 3 R71 af T 6 Cag t l rgo e lgo t lao u where ie the element resistance at T C R is the element resistance at O C and 2 amp and 6 afe constants for each individual platfnum element A platinum probe system provides greater output voltage therefore has greater tolerance to noise than thermocouples and does not require a reference junction temperature or other compensating device The platinum probe is superior to most other methods of on board temperature measurements but care must be taken in signal conditioning to insure that effects such as self hcating do not occur Signal conditioners specifically designed for platinum probes are available and can produce excellent results For instrumentation systems requiring both cockpit display and data recording dual element probes are available to prevent undesirable interaction of electronics 2 1 2 Atitude Altitude measurement is accomplished in terms of atmospheric pressure and height above ihe grourd The data is required in the cockpit so the pilot can stabilize at an altitude or maintain a prescribed path relative to the ground The information is also made available to the instrumentation system Typically pressure altitude measurements will range from 60 m 200 ft below sea level to 7500 m 25 000 ft dur
18. position transducer Lag angle is also measured at tre blade root in a similar manner These transducers are generally potentiometers or linear variable differential transformers Blade azimuth during each revolution must be measured to evaluate significance of blade behavior The main rotor speed can be measured with a tachometer however this does not give the azimuth of a particular blade at a point in time One way to record blade azimuth is with stationary receivers which sense passage of a magnetic or optical device attached to the blade The number of sensors needed per revolution will depend on the accuracy with which the azimuth must be established The sensor signals can also be correlated with the rotational speed measurement from the tachometer Optical and acoustic devices may also be used for blade position measurements 5 2 Hubs The rotor hub experiences large tension loads caused by centrifugal forces on the blade bending moments from in plane motion vertical moments from blade flapping and torsion caused by blade pitching moments Rotor hubs are usually complex forms and stress analysis is required to determine the location for the strain gages The number and location of the sensors will be unique to each installation and must be established on a case by eee basis Resistance strain gages are sed in a manner similar to that discussed for the blades 5 3 Pitch Links The collective and cyclic control is transmitted to the blad
19. radar altimeter is used to supplement the pressure al l cayrrement The antenna for the radar altimeter is mounted to insure no return s h fr the airframe This is a particular problem for aircraft with fixed gear or wrid Tr cockpit indication of radar altitude is of special importance during tests near the vcouud and n some cases a 3m 1 ft resolution is required In all cases the pilot ta w uno eight within 3 m 10 ft A typical radar altimeter is the Honeywell model AP s1 Inia altimeter nas an accuracy of 5 m 1 5 ft plus one percent plus five per ut of he average range rate and ofvers a test mode switch for system checkout and preilight Auxillary outputs are used on the radar altimeters to provide inputs to the ins umentation system for recording both absolute height and rate of change of height aese signals are most often in analog format 2 1 3 Humidit At a given atmospheric pressure and temporature humidity can affect helicopter performence relative to dry air by a decrease in power available or an increase in power required In the first instance the humidity decreases air density and thus mass flow through the engine and in the second case the rotor will experience an effective in erease in density altitude The effect of humidity can cause several percent error in the density which can have a significant impact on helicopter performance Humidity effect while large in theory have not adequately been measure
20. resistance to flow which may affect engine operation and should be considered prior to installation The sensor is placed in the fuel line as near the engine as practical Extreme care must be used to insure that the installation plumbing is the same as that used during the calibra tion Variations can cause flow effects which will invalidate the calibration The speci fied flowmeter installation requirements should be met These usually specify the diameter of the input output plumbing length of straight line tubing required at input output and may provide limits of vibration exposure 3 7 2 Fuel Temperature f To obtain a vrue mass fiow the fuel temperature must be measured at the same place as the volume flow was measured When the measurements are made very near the 3 engine the system will usually be enclosed by the cowling and fuel temperatures up to 50 C are not uncommon Consideration must be made for any oil fuel heat exchanges in the system Fuel flow measurements are usually made at relatively steady flow conditions and for maneuver situations dynamic respon s of the sensor must be considered The tempera ture sensor is pluced directly upstrean of the flowmeter A suitable type of sensor for Document provided by SpaceAge La are Can ated itd haa Ris dnw Aau et setae Sle Control Inc ha ite http spaceagecontrol com Savini sek meta a LEPE ris F this application would be a platin
21. speeds Transfer of information from rotating to stationary members is a particular problem In addition rotating parts experience various types of loads and quickly amass a high cyclic count Testing must be accomplished to establish a fatigue life for each part Vibrations in all directions are prevalent and encompass a wide range of frequencies and amplitudes The aerodynamics of the helicopter produce stability and control characteristics that often require improvement through use of mechanical hydraulic or electronic systems Complex control systems require that actuator motions and electronic inputs be measured Instru mentation must consider each system relative to the basic flight controls Rotor blade information can include static and dynamic pressures positions angles and stresses The blade tip may be in the transonic regime while reverse flow may exist at the hub Blade instrumentation must have minimum influence on the lift and drag characteristics The large volume of air displaced by the rotor at hover ani low speed can have a strong influence on weapons firing or personnel working near the helicopter In this flight regime the helicopter is usually near the ground and the downwash can introduce environ mental problems associated with high velocity and hot engine exhaust gases Combined rotor wash and flight in any direction may require measurement of unusually large angle airflows into the engine inlets and on lifting surfaces The r
22. taken into consideration when selecting suitable sensors 4 1 1 Pitch and Roll Attitude Attitudes are measured with a 2 axes gyroscope For most engineering flight tests a nominal range is 45 in pitch and 60 in roll with a 0 5 resolution The resolution is dependent on the signal conditioners and encoders Electrical conversion of pitch and roll position is accomplished in most instances by one of two methods Synchro signals are in some cases used for position information output but this method requires conver sion of the synchro data into an electrical form compatible with the data recording systems Document provided by SpaceAge Control Inc meer x Se og E TE Bes Seite REEE a La sizaki Lied deeds a bitten oc ee heh lee i http spaceagecontrol com 0 o LE PINE RESSE ATTON NETIC EY BOE OD ACR A PCT gg signal conditioners or encoder More direct input of position data to the recording system can be achieved by using gyros with potentiometers for position encoding These are most often directly compatible with recording system signal conditioners The attitudes are measured for steady state conditions as well as deviations from the trim during maneuvers 4 1 2 Yaw Attitude Yaw attitude must be measured from some reference point The gyroscope is similar to those for pitch and roll with the addition of a caging feature At trim or some desired J starting condition the gyroscope i
23. the YC 15 Advanced Medium STOL Transport March 1977 AFFTC TR 7641 32 H K Cheney YO 15 STOL Performance Flight Test Methods Eighth Annual Symposium Proceedings of the Society of Flight Test Engineers 33 B K Parks Flight Test Measurement of Ground Effect Eighth Annual Symposium Pro ceedings of the Society of Flight Test Engineers 34 W C Bowers R V Miller Inertially Derived Flying Qualities and Performance Parameters Society of Experimental Test Pilots Symposium Proceedings Sep 22 25 1976 35 J N Olhausen Jr The Use of a Navigation Platform for Performance Flight testing Society of Fiight Test Engineers Symposium Proceedings Aug 21 23 1973 Document provided by SpaceAge Control Inc http spaceagecontrol com ERRETEN eee tif eet PRSE Se see Gi et hag SH PERSER SSE betes arran WS ASIP EDT thee ET Ao rman eo PRs NEI gh mute TEEN e do DRE a ee At PU ERE L TOS YT AE TTT ES FETTER RLY NERV TT J APPENDIX I TYPICAL HELICOPTER INSTRUMENTATION REQUIREMENTS This appendix provides typical requirements for a helicopter instrumentation installation While certain characteristics can be specified for each parameter it is usually necessary to make adjustments dependent upon the nature of the test Some common variances have been noted in the remarks section The accuracy stated is base
24. the helicopter flight envelope and during the mission requirements The data is obtained during ground tests Climatic hangar tests and flight tests In addition to vibration which has previously been discussed the compartment temperature quality of air and noise environment are of primary concern Ground tests and climatic hangar tests often generate absolute data which is used ir assessment of basic design or hardware modifications Flight test data is usually evaluated in terms of how the conditions affect the occupants SDP LS PE ROSS oe 4 7 1 Air Tomperature and Airflow The stabilized temperature within the compartment is dependent upon the outside ambient conditions the quantity of heating or cooling added the efficiency of the distri bution and the heat transferred from the compartment to the outside The vertical and lateral temperature gradients should be measured at the crew stations Solar radiation or extraneous heat sources should be considered when selecting sensors Thermocouples shielded or unshielded as required provide satisfactory results Depending on the accuracy desired and temperatures to be measured the systems in use range from iron constantan thermocouples to platinum element probes with very exacting wiring practices and signal conditioning The number of sensors and the distribution within the space can be based on a human factors evaluation analysis of the airflow pattern or qualitative judgemen
25. H4 and requires J E excitation Careful selection of pre sampling or signal conditioner low pasa filtering a will eliminate the acceleration inputs from the sensor above a predetermined maximum frequency of interest The general frequency band desired is fron D C to main rotor frequency Selection of the vroper accelerometer should consider ambient conditions size constraints the acceleration range present and the frequency range desired As with attitude gyros the accelerometer orientation should be exactly defined by the aircraft axis Accelerometers are usually calibrated to the standard gravitational acceleration value For some tests output may be corrected for local gravity Te Fie al a 4 5 Vibration Airframe vibration frequencies are predominantly multiples of the main and tail rotor speeds An out of balance or out of trim blade will generate a vibration with a frequency equal to the rotor speed Helicopter main rotor speeds may produce vibration fea RS nied Beth AACR NT Arent cent eni A SRE ies at oer Bs teeta Document provided by SpaceAge Control Inc http spaceagecontrol com See pA HELLS FP TIE LRT PRD TERT eer roe TEST TS Se as m iti a frequency as lew ds 3 HZ Other vibration sources usually have frequencies greater than those from the main rotor It is necessary to define both frequency and amplitude in order to evaluate the effect on structures components and occupants The testing
26. J pa NORTH ATLANTIC TREATY ORGANIZATION ADVISORY GROUP FOR AEROSPACE RESEARCH AND DEVELOPMENT ORGANISATION DU TRAITE DE L ATLANTIQUE NORD iei a oo E Volu 1g HELICOFT ERI F MOGHE TESTI INSTRUMENT ATION BEFASI hy QF nne h R Ferrell i x Volume 10 t of the AGARD FLIGHT TEST INSTRUMENTATION SERIES Edited by A Pool arid K C Sanderson an i 4 4 a P 4 a 4 are a ea Dt wn Sees SE Sue ig This AGARDograph has been sponsored by the Flight Mechanics Panel of AGARD _Dogument provided by Spa Age Control Inc Yip gt Spacga gs sont C7 aay ago sa ainor AVAA NO AA S Mas ATTERT an SGAN i TOT TPIT HUTT PRS ATE TOD FAERIT brine 9 SUMMARY 1 on i 1 INTRODUCTION 1 1 Test Vehicles 1 i 1 2 Types of Teats 2 B 1 3 Instrumentation Environment 3 1 4 Systems Planning s ih 1 5 Installation 4 7 2 ATMOSPHERIC i 9 H E Y 2 1 Air Data Instrumentation 4 te 2 1 Free Air Temperature 4 a 2 1 Altitude 5 p 2 1 3 Humidity 5 i 2 1 4 Icing 6 i 2 2 Airspeed and Relative Wind Direotion 7 K 2 1 2 Swivel Head Test Systems 7 ia 2 2 2 Umani Directional Airspeed Systems 8 i j a Aeroflex 8 e b Elliott 9 fh c J TEC 10 if j d Loras 11 i e Rosemount 12 Ke t f Honeywell 13 i i 3 PROPULSION SYS EM 14 3 1 Shaft Speed Measurements 14 A 3 1 1 Engine Speed 14 3 1 3 Drive shaft Speed 15 3
27. Less noise and better quality data c Less maintenance and more reliability During development testing testing of new rotor systems or flight in extreme aerodynamic conditions where rotor behavior must be closely monitored a telemetering system on the rotor hub may be used with recording and analysis equipment on the ground Document provided by SpaceAge Control Inc http spaceagecontrol com aia Maitai i a ea ai aS Pit aa ee eai irh pau 5 5 Non Rotating Surfaces ra ata RRA aiaa AFA aa a a a S a Many helicopters use horizontal and vertical stabilizers for stability and con trol and wing lift to augment rotor lift The stabilizers can be fixed connected to the flight controls driven by electronic controls systems or dynamic pressure The operation of the system will determine the instrumentation required instrumentation for stress and pressure distribution is accomplished as discussed for the blades Surface position is usually measured as control and displacement or angular deflection from a specified zero or trim point Transducers for those applications are usually potentiometers or linear variable displacement transformers 6 FLIGHT CONTROL SYSTEM Helicopter flight control systems vary from the most elementary mechanical arrangements to very complex systems which include mechanical hydraulic and electronic components Control inputs in the cockpit are transferred through the system to the rotor
28. Sensitivity and response are the most important criteria in selecting a sensor which will meet the data requirements Time averaged data suitable for most requirements is provided by a low response sensor such as a resistance probe Small rapid variations such as required for quality of airflow are obtained with hot wire or hot film anemometers i Tea 8 1 4 Ambient Air Pressure The ambient air pressure is usually measured at ground level and a standard lapse rene for height above the ground is subtracted to obtain the pressure at the test venicile 8 2 External Noise The helicopter noise is generated by the engines transmissions and rotors The helicopter internal noise measurements were discussed in section 4 7 3 The external noise is measured with fixed or portable ground located sensors The noise is evaluated for tactical suitability for the military and in terms of environmental impact when opera ting in the public sector Helicopters are inherently hoisy despite efforts to reduce noise levels These efforts include tranemission design and manufacture noise isolation techuiques insulation application as well as number of blades and airfoil design The sensor performance will be influenced by the atmospheric environment and these parameters must be measured in addition the aabient noise level must be established prior to taking measurements of the test vehicle noise 4 k Nn aa ote i et aa OS co te ee ee Docu
29. YSTEMS 7 1 Forcea and Motions 7 2 Firing Effects 7 3 Ejected Material 7 4 Fire Control System 7 5 External Noise 7 6 Gas Contamination 8 GROUND SUPPORT INSTRUMENTATION 8 1 Atmosphere 8 1 1 Wind Speed 8 1 2 Wind Direction 8 1 3 Ambient Air Temperature 8 1 4 Ambient Air Pressure 8 2 External Noise 8 3 Thrust 8 4 Space Positioning 8 4 1 Instrumented Range Operations 8 4 2 Remote Site Operations 9 REFERENCES Appendix I Typical Helicopter Instrumentation Requirements Appendix IT Typical Instructions for Developing and Maintaining Recorded Instrumentation Parameter List http spaceagecontrol com GRABIVIESUUA Jtron j Al A2 we mbit Ae was Fs Baa Se H Send 2 ae hanian BT ah AD ety S pagina T 5 i A N X EEE oe a a ee ee eee RUMENTATION Kenneth R Ferrell US Army Aviation Engineering Flight Activity US Army Aviation Research and Development Command Edwards Air Force Base California 93523 HELICOPTER FLIGHT TEST INST SUMMARY This document discusses the helicopter characteristics with which the instrumen tation must contend and outlines typical tests that are conducted Major aircraft com ponents and systems which may be instrumented are listed and suggestions are made for sensors locations and installation Details are provided for instruments peculiar to helicopters Interface of the test instrum
30. a ture of the surface should be used as one criteria for selection of the glue Further comment on accelerometer useage will be given in the discussion of airframe vibration Velocity transducers are generally of two types piezoelectric or a permanent magnet coil combination The piezoelectric transducer is actually a piezoelectric accel erometer with an integral amplifier integrator and has a frequency response range of approximately 1 HZ to 2000 HZ The permanent magnet coil combination transducer has a lesser frequency range typically 10 HZ to 1000 HZ but is self generating and therefore does not require a regulated power supply as does the piezoelectric transducer Both provide a millivolt output proportional to velocity 4 AIRFRAME The airframe measurements are needed for a variety of tests Performance testing includes drag determination which requires attitude and relative wind data In some cases the data must be corrected for linear or angular accelerations Attitudes rates and accelerations are critical for stability and control tests In these tests the struc tural and dynamic maneuvers require measurement of loads or stresses in various components to establish component life of flight envelope limitations Airframe vibration information is needed to evaluate occupant environment and conditions experienced by instruments and aircraft sub systems 4 1 Attitude The aircraft attitude is measured relative to an earth axes syste
31. ability to change height above the ground and vary heading through 360 Ref 24 Another system involves suspending the aircraft and sensors are used to measure the changes in the forces Ref 26 The sensors are usually an integral part of whe thrust stand and the data is recorded on the ground Since the aircraft is restrained the easiest way to transfer data is by electric cabling For other than thrust stand operations a suitable ground restraint system with great flexibility can be constructed with instrumented cargo hooks or load cells Load cells may be constructed by the test instrumentation group or a commercial sensor may be used The commercial equipment comes with various ranges and the sensor selected should be compatible with the expected thrust of the particular helicopter The load cell is placed in series with cables attached to a ground restraint It is important that the inatal lation does not allow the load cell to drop and be damaged when the emergency cable release mechaniam is activated When the load cell is ground restrained an electrical quick release must be placed between the aircraft and the load cell The sensors are usually strain gages and the output is wired into the airborne data system and when possibile is ground recorded The longitudinal and lateral deviation angles are measured with iinear accelero meters mounted on the load cell When vertical the accelerometers read zero G and when horizontal the outp
32. aceagecontrol com RRR aE ae hs ieee E E A P eee PE TAR E OMAHEN IRANOM fest AIEA RN aai etA edith 7 Copies of the typed Recorded Instrumentation Parameter List are distributed after the last calibration The master list is filed in the aircraft instrumentation file 8 Normally six 6 copies are made and distributed two to the F T E one in the aircraft one in the Data Processingfile one to the instrumentation technician and one retained by the data systems technician as the master correction draft The F T E may specify other distribution 9 All copies will be made in reduced size 8 x 11 and on both sides if two sheets are required so that the complete information is on one piece of paper Normally a few minor changes can be made by hand on the copies After several changes have accumulated the list will be revised using the following procedure 1 Last effective flight for that list version will be filled in and a full size copy made and filed for historical purposes 2 New applicable first flight last change date and date flight and reason for change of each affected parameter will be completed and retyped 3 Copies and distribution will be made as on the initial list 4 If applicable a new calibration deck listing will be made copies dis tributed and filed There should be a direct correlation between the Recorded Instrument Parameter List versions and calibrat
33. acle shipboard landings or autorotational landings may generate sigzificant normal acceleration loads 1 4 Systems Planning The instrumentation system must be carefully planned to insure that the necessary data will be recorded in the best manner within the physical and cost limitations The recorded data may be used in different forms or may be processed in several ways which require consideration of the data processing facility The test objectives must be care fully analyzed to determine the required number of measurements These results deterzine the size of the installation and the recording device They also have an impact on the method of recording Volume or weight conflicts may arise which dictate priorities among the desired information Essential desirable and non essential items can then be deter mined accordingly Test requirementa set the initiai data accuracy goal and then appropriate system characteristics are established Required data accuracy must be considered for all system components Transducer requirements are established and signal conditioning is designed Throughout this effort the magnitude of the expected error must be known Close coordina tion must be maintained between flight test and instrumentation engineers to insure that accuracy requirements are not overly stringent Compromise or relaxation of the require ments may be needed to prevent escalating complexity or cost Helicopter flight tests usually vequi
34. amp a AIRCRAFT ANGLE OF ATTACK AB 2 AIRCRAFT ANGLE OF SIDESLIP WsWIND VECTOR K Figure 2 2 2 6 Honeywell Airspeed Sensor The system is a sensor head and an associated electronica package The sensor head has three receiver probes spaced at 120 intervals around a transmitter and a tem perature sensor The transmitter is a piezoelectric transducer resonant at 75 kilohertz and the receivers are wide band width ceramic nicrophones with response to 400 kilohertz The temporature sensor is a platinum element thermally isolated from the structure The sensor unit also contains a temperature sensor amplifier and three receiver preamplifiers The transmitter drive timing logic pulse detection circuitry and electronics used to solve the equations are contained in the electronics unit Document provided by SpaceAge Control Inc ORN http spaceagecontrol com tataia coco dell feasts MAME hiia ris lav cela i ie pe A ERY Ty iat ot gt ere ass wheres mio eet Ses ag se 3 lie Sat dad ii ii Pebining the wind velocity components is done with a geometric arrangement of three ultrasonic transmission paths deployed in the airflow From this three equations can be derived to express the velocity components as functions of the measured trans mission times slong the paths The temperature sensor is needed to compute the wave velocity in air as a function of
35. apons Firing S Stores Jettison Envelope i Instrument Flight Capability Aircraft Systems Failures v Simulated Engine Failure 4 i Automatic Flight Control System Failure Hydraulic System Failure Tail Rotor Failure Autorotational Entries Autorotational Landings hw i Structural Dynamics fs Vibration E Structural s E Human Factors Cockpit Evaluation Night Evaluation Internal Noise Temperature Toxicity i Reliability and Maintainability Subsystem Tests 1 Engine Performance x Aircraft Pitot Static System i Weapons System Electronic Equipment and Antennas Hydraulics Environmental Aspects External Noise Radar Reflectivity Infra Red Radiation Downwash Effects Types of instruments ranges accuracies and environmental aspects must all be considered The optimum situation is to have a fully instrumented aircraft capable of recording all parameters However for some tests satisfactory results can b obtained with limited instruments at considerable time and cost savings The most exacting instru mentation requirements are for the performance tests In these tests quantitative data are the primary results and subjective opinions are used to evaluate pilot ability und machine capability relationships Power measurement is the most difficult and most important Small helicopters often have engines in the range of 150 to 225 KW 200 to 2 50 UHP and a one percent error is most difficult to measure A limited amount of a
36. ateral and vertical components A resultant is not presented nor is angle of attack or angle of sideslip calculated Free air temperature is measured and the computer calculatrs true airspeed Individual longitudinal and lateral airspeed indicators type 71 011 01 consist of a stepper motor and a feedback potentiometer This provides an indicator rate signal and position signal which is fed back to the airspeed computer The signals are summed with the computer longitudinal airspeed and are checked by the servo monitor Detected failures are indicated by a warning flag on the indicator ec J TEC The VT 1003 vector airspeed sensing system is manufactured by J TEC Associates Inc of Cedar Rapids Iowa The J TEC vector airspeed sensing system measures relative wind speed and direction with no moving parts The VT 1003 consists of a sensor head an electronic processor and an airspeed and direction indicator The sensor is illustrated in Figure 2 2 2 3 The sensor head consists of six identical tubes 6 67 cm 2 5 8 in long mounted radially on a 13 65 cm 5 3 8 in diameter hub It is mounted on the aircraft so that one pair of tubes is aligned with the lateral axis of the aircraft and the other tubes are 30 degrees either side of the longitudinal axis The sensor weighs approximately 1 6 kg 3 1 2 1b Regardless of wind direction flow exists in at least two adjacent tubes at any time allowing two equations to be solved simultaneously
37. blank are detailed below PROJECT Project Number AIRCRAFT Test aircraft designation status mission type model and series S N Serial number tail number or other designation Note Programs calibration decks apes etc will be filed by 8 N SHEET Number in sequence of sheets required for complete list normally two will be uar i e 1 of 2 and 2 of 2 EFFECTIVE FLIGHTS THROUGH Sertes of flights for which this version of list is applicable i e ghts ru 149 Last effective flight will be filled in when list is updated and before new copies are made ORIGINAL LIST DATE Date of completion of original basic list LAST CHANGE DATE Date when list was updated for recent changes and new copies were made and distributed COLUMN ENTRIES PARAMETER Measurement name DATA PROCESSING E Mnemonic name assigned to the measurement in data reduction software Document by TES Control Inc htt bi spacea econtrol ec y aa ats a FL SRS FRSE rd tt 4 4 H nh PRG TITS igital Signals example First two characters PD Parallel digital SD Serial digital Second two characters NB j Natural Binary BC Binary coded decimal 20 2 s complement 3C 3 s complement OB Offset Binary GR Grey code Third two characters Number of bits decades or octaves LOCATION Location of transducer measurement usin
38. caused by tying signal conditioners to the control signals In control system development it may be necessary to instrument for hydraulic fluid pressure or flow 6 4 Swash Plate The resultant of all control inputs culminate at the non rotating swash plate where they are transferred to the rotary control system In certain stability and control analysis the swash plate angle is a requirement The longitudinal and lateral angles relative to the shaft or the airframe are measured with position transducers 6 5 Stability Augmentation Systemss Most helicopters have systema to improve the stability and control characteristics Increasing use is algo being made of flight director systems which assist the pilot in normal flight and increase the capability to operate in adverse weather conditions These systems may be self contained or may use parts of the standard helicopter systems The systems are essentially a computing device which receives information and on the basis of calculations or predetermined logic makes an input to the flight control system The instrumentation needed for the systom varies greatly with type of test being conducted During control system development and optimisation tests each input and response must be measured For conventional flight tests the system input to the flight control system is Document provided by SpaceAge Control Inc htto spac A A Mires NC ND Sk age panasan mesan t eate Aiai a AALL AS 4 4
39. ce contact results from improper mounting Alignment of the sensors is critical and should be traceable to a known reference 4 5 2 Sensors Velocity pickups or accelerometers are suitable with the latter being in most common usage It may be necessary to sense vibration in a single axis or in 3 axes A suitable single axis sensor is a piezoelectric with a frequency response of 5 vo 2000 HZ Piezoelectric accelerometers are a good choice because they have self generating output wide frequency response small size and are easily mounted For dynamic acceleration which is of interest in airframe assessment the piezoelectric sensor also has the advan tage of not responding to input frequencies much below 3 HZ These devices do require some care in application however By employing the piezoelectric effect the sensor produces a charge that is proportional to the acceleration level It is then necessary to convert tais charge to a voltage for input to the instrumentation signal conditioning The voltage produced will be proporticnal to the charge and the capacitance of the sensor cable and signal conditioner input combination Unfortunately the capacitance of the cable will vary with length which hampers interchangeability and by flexing the cable charge noise can be generated and is indistinguishable from the sensor charge output These drawbacks can be overcome by using a well placed charge amplifier rather than a voltage or source follower amp
40. ckpit indicator for use in establishing flight test cunditions This indicator should have at least 1 C insrements Helicopter flight test temperature conditions may vary from climatic hangar or arctic testa at 55 C to a desert condition where the temperature is 55 C For other than extrene environmental tests a commonly used instrumentation range is from 35 C to 50 C A platisum element resistance probe is generally used to sense the free air temperature Pure platinum has been selected as the international standard temperature measurement from 182 97 C to 630 5 C and wher properly used and calibrated accuracy to 0 1 C can be realized in field operation To achieve accuracies of this magnitude care must be Document provided by SpaceAge Control Inc http spaceagecontrol com ARES ti te intatt Misti teuria aiii aE rioarekin W i E gt 5 ead ta sop ba aici SS a EE reer AEA TROTA STR ZR a eet PN TEAL TLE e ep AN n EE A BATSERIO EE e ip ere a ea setae a aa cae COAL Wien SENDENTA EALA AAA at aaa oo odd iait taken in both calibration and system integration of the probe For a calibration covering the entire range of the platinum probe measurements of probe resistance are made at four specific temperatures and these values are employed to generate values of resistance for any
41. cording optical tracking instruments use a telescope to track the helicopter and record the data on film These systems have dif ferent accuracies and capabilivies A principal instrumentation consideration is that the data recorded outside the aircraft is difficuit to correlate or merge with airborne recordings A typical solution is to photograwh external event lighte on the aircraft while recording time or electronic identification data at other locations In addition the sample rates are inadequate to obtain accurate acceleration data The radar range systems may use either pulse or continuous wave cw equipment The most frequently used system is a pulse type with high peak power wide band width signal transmission and a highly directive electro mechanically steered antenna The apparent radar range is derived from the time needed for the pulse to reach the target and return Tracking system electronics maintain the antenna parallel to the returning wave front and the bearing is measured by tracking system The data output is range angles and rate of change The operation of the system and the many corrections which must be applied to the raw data are usually beyond the capability of the flight test personnel and must be accomplished within the range facility Various equipment or tracking problems can be expected because of low angle multi path and refractiun conditions high target acceler ations and target radar geometry The cw radar determ
42. ctively when the airspeed sensor is mounted parallel to the aircraft s vertical axis Those indicator limits were chosen to provide maximum sensitivity while encompassing the expected range of helicopter operation PRE TE Pe SENT OME SE The indicator scale is presented in the form of concentric rings located at 10 Kn 5 m s circle increments with zero located at the geometric center of the indicator and a 40 Kn 20 m s circle being the most distant ring The horizontal pointer reflects forward velocity by moving upward and rearward by moving downward The vertical pointer indicates transverse velocity right left A left vertical pointer deflection indicates flow coming from the left and similarly for right deflection a flow from the right Viewing the intersection of the horizontal and vertical pointers will depict the vector resultant of airspeed f Honeywell The Ultrasonic Wind Vector Sensor UWVS is designed and manufactured by the Government and Aeronautical Products Division of Honeywell Inc St Louis Park Minnesota The system was designed to provide an accurate measure of the relative wind while using no moving parts giving linear sensitivity over the airspeed range and responding to rapid changes in wind magnitude and direction The UWVS operates on a principle involving ultra sonic signal transmissions through the moving air mass The sensor and relative wind vectors are shown in Figure 2 2 2 6
43. cument provided by kamak Control Inc http spaceagecontrol com saranin Nia ot eaten eneiniad Be ada estas a de Spat Ae Np ese ah oe wae j 4 1 ee aes Ee iF Alcea theca ett E EE baka EER ET ER j Instrumenting for a detailed engine temperature evaluation is an extremely diffi cult task and should only be conducted with the assistance of the engine manufacturer This is advisable because of potential engine performance changes or structural impli cati ns caused by the test installation The temperature evaluation may be for internal gas flow parameters or for external surface and compartment conditions The internal engine temperatures are ambient at the inlet and increase to a maximum in the burner section Engine design information can be used to select sensors which are suitable for the range and response requirements Most flight tests require measurements of only compressor discharge power turbine and exhaust gas temperatures These measurements are commonly taken from standard engine sensors which are usually available Additional temperature information is obtained by placing thermocouples at selected stations shown in Figure 3 5 1 Engine surface and compartment temperature measurements are necessary to estab lish the engine environment and assess the heat being transmitted to the surrounding structure Surface temperatures can be measured within t1 by use of color coded tempera
44. cumentation of ice accretion The buildup on the rod gives an indication of ice accretion on non aerodynamic surfaces while the airfoil is of conditions on lifting surfaces and may correlate with main or tail rotor conditions The Rosemount ice detector uses magnetostriction to drive a sensing prove at its naturel frequency As the probe accretes ice the natural frequency changes due to the increased mass The change is calibrated in terms of ice accretion rate The calibration of such a system must teke into consideration factors such as airspeed which affect ice accretion When the ice thickness reaches a predetermined value the probe is deiced and the cycle repeated Cycle counting can be used to obtain total ice accretion The probe is housed in an electrically heated aspirator shroud which uses engine bleed air to induce ambient airflow over the probe during hover and low airspeed he Leigh ice detector consists of a light emitting diode photo transistor assembly which provides an optical path that is partially occluded by accretion of ice on the ice detector probe The assembly is encased in an annular duct and ejector nozzle which is supplied with bleed air to induce high velocity airflow over the ice collecting probe and provide anti icing When the ice accumulation reaches a pre set level the probe is electrically deiced and the cycle is repeated The icing signal is displayed on cockpit indicators and recorded by the data syst
45. d in flight However measurements ot humidity should be made in order to build up a data bank for further analysis The criticality of density changes increase with higher temperature and higher relative humidity The density can be measured directly with nuclear radiation devices Refs 2 end 3 The accuracy of the referenced devices are 1 to 2 as they existed at the time Increased accuracy can be obtained by increasing the signal strength However extreme care must be used relative to the radiution hazards Electronic hygrometer equipment is also available to measure the relative humidity directly Quoted accuracy is 1 5 When engine power is being corrected for humidity the measurement must be recorded for each test condition Independent measurements of free air temperature and dew point allow calculation of relative humidity and the effect of air density Refs 4 and 5 Document provided by SpaceAge Control Inc http spaceagecontrol com ete GRE ets eH Pa ei Nan a at i Ta NUT TREE ATEN EEU EEE FURER AN AP ETN BPAY EA ERRA four IE i 2 1 4 Icing Helicopter icing tests require that the water characteristics of the cloud be s measured for correlation with ice accretion and effects on the performance or handling l qualities of the helicopter Measurements include droplet size and distribution as well as liguid water content The airflow character
46. d on the engine to measure motion in all axes The sensors must be located relative to any absorbers or dampers so that the desired vibration is being measured In the case of more than one engine each must be instrumented because of potentially large changes with asymmetric power The sensors are usually placed on the engine mounts Transducer types employed can vary greatly due not only to environmental and parameter requirements but also as a result of the analysis philosophy empioyed In most instances accelerometers are used to assess vibration levels but some analysts use velocity sensors instead No attempt will be made in this document to influence the reader in favor of one method or the other but a brief discussion of transducer types is pre sented Vibration levels at engine stations are due to a broad frequency spectrum of iuput The rotor system excites the area with low frequency while rapidly rotating de vices including the engine provide high frequency excitation Most often the full range of inputs can be sensed using piezoelectric accelerometers These have a flat frequency response of approximately 3 to 30 000 HZ Mounting of the accelerometer should be accom plished without changing the vibrational characteristics of the test article That is the mass of the accelerometer and the mounting device or material should be carefully chosen Often glue is used to attach the accelerometer If this method is employed the temper
47. d primarily on the data requirements and resolution should be adjusted as necessary GREAT CARE SHOULD BE USED TO SPECIFY NO MORE ACCURACY OR RESOLUTION THAN IS ESSENTIAL Conversely instru mentation should comply with the specification easible and of course any additional capability will enhance the quality of the results The helicopter can be expected to generate vibrations which will affect each sensor used The vibration frequency will vary from a one per main rotor revolution to high speed jet engine frequencies Each sensor should be evaluated with respect to the driving frequency it should experience For a particular location this value is then used to establish the maximum frequency of interest for electronic filtering and signal conditioning In many cases a careful study of the sensor characteristics will greatly reduce the amount of electronics needed REFER RESO Ter 1 2 1 2 Pressure 1000 to 5 ft Altitude 20 000 at SL 2 1 2 Radar 0 to 1000 Altitude 2 1 2 Pressure Rate ft min of Climb Radar Rate ft sec 100 1 0 1 0 of Climb PARAMETER 2 1 1 Free Air Expand Range to 85 C for coll weather testing Quartz capsule with microcomputer 100 ft vernier with 5 ft increments for cockpit Temp3rature Dew Point Temperature Liquid Water O to 3 Content Pitot static Airspeed 2 2 1 20 to 250 1 0 or Swivel head with 20 0 freedom in ail directions oe Pee 1
48. e With a large number of sensors it may be necessary to use a time dependent sampling technique A large time increment between samples will restrict capability to establish variations in the temperature and a continuous record of selected sensors may be necessary 3 4 3 inlet Devices The status of inlet devices such as guide vanes by pass doors or variable geo metry equipment must be recorded This information is needed to evaluate and corkelate inlet flow characteristics and calculate airflow In addition there may be drag consider ations These inlet devices are usually mechanical and position sensors such as potentio meters or micro switches are used to record their motion There may also be inlet ram air bleed devices and it may be necessary to measure the flow taken from the inlet 3 5 Engine Temperature Engine temperature requirements can vary from a single parameter to detailed measurements at various engine stations Engine stations and nomenclature usually varies with each engine however the system to be used must be defined prior to instrumentation system design A typical engine layout and definition is shown in Figure 3 5 1 STATION NUMBER 9 pe 10 gt 7 gt gt gt gt FREE FREE STREAM EXHAUST AIRFRAME AIRFRAME INLET EXHAUST ENGINE LAST INLET TURBINE AND FIRST ENGINE EXHAUST COMPRESSCA FIRST Assy TURBINE ASSY COMPRESSOR SUMER DIFFUSER Figure 3 5 1 Engine Station Definition Do
49. e cockpit control is usually a lever and from the mechanical advantage and the measured force the actual pilot input force can be calculated Cockpit control motions are measured with position transducers Those trans ducers are placed on the control rods attached to the cockpit controls The transducer output is calibrated by moving the controls through fuli range of travel 6 2 Mechanical Linkages In structural tests it may be necessary to instrument all linkages from the cockpit controls to the stationary swash plate This instrumentation will usually be strain gages or position transducers which are treated as discussed under cockpit controls 6 3 Actuators Control systems may have hydraulic or electrical actuators which transfer pilot inputs to the rotor The actuators may also be driven by any stability and control devices The resultant inputs are thus a sum of all inputs and must be measured to obtain the net control input Hlectrical inputs are measured and the voltages are then used to determine equivalent linear deflection of the control input The actuator stroke may be measured or more commonly the motion of the control member connected to the actuator is measured These motions are sensed with a position transducer Any electrical signals within the control system must be measured in a way which insures no change in the signal character istics This may require an isolation type of amplifier to eliminate any detrimental effect
50. e collection efficiency cf each cylinder is different and thus accretes ice from different droplet sizes From the amount of ice on the different cylinders a pro AEL M of droplet size and distribution can be constructed For other than conditions of o K temperature and low liquid water content the cylinders have limitations which Can A a Significant errors Refs 7 and 8 A newer method uses a laser ager lent wii Knollen berg probe This instrument operates on the principle that the 1 light will be n scattered by the droplets as they pass the light J 1 system collects the i scattered ligbt and through electronic means the pract size and distribution is Cetermined Refs 9 10 and 11 Each probe is designed for a range of droplet sizes and di care must be taken to insure that ee uani is used to encompass all the droplet d sizes The output from the laser sys an be recorded on magnetic tape or with proper equipment can be viewed in real time l Liquid water content can be calculated from accreted ice or measured directly in i 5 the atmosphere Ref 12 The previously discussed rotating cylinders accrete ice which can x be removed and in conjunction with the collection efficiency can be used to calculate the t liquid water content The visual ice detector probe has a small airfoil with a steel rod l protruding forward of the leading edge The protruding rod is marked or color coded in A increments for visual or photographic do
51. e is then determined A transit is placed at the end of the center line projection and offset laterally the same as the boom mount ie from the aircraft center line By sighting the length of the boom through the transit adjustments are made to the boom until it is parallel to the line of sight This insures the boom is aligned with the aircraft center line A similar method is used with a plane through a water line to assure proper angle of attack alignment Free stream angle of attack and sideslip are obtained by mounting movable vanes on the airspeed boom A vertical vane gives sideslip and a horizontal vane provides angle of attack Suitable fixtures are constructed to calibrate the vanes relative to the boom A typical swivel head probe and vane installa tion is shown in Figure 2 2 1 1 rren a E akian i a ES a aes ee Sg PITOT STATIC SWIVEL HEAD PROBE Sa Eee SF i et a a eT a an tle a lt gt sae gens 21 MINIMUM ANY DIRECTION Srp NY PITCH WANE Ya VANE Figure 2 2 1 1 Boom Installation of Airspeed Angle of Attack and Sideslip System iaki A suitable pitot static probe has a swivel head gimbal freedom of at least 20 degrees in all directions airspeed capability up to 100 m s 200 Kn angle of attack and sideslip motion of 45 degrees Airspeed error should be less than 0 5 at angles of attack and sideslip to 30 dagrees The vanes are mass balanced and airfree to alian with the relative airfl
52. e the aircraft is moving within the air mass Atmospheric data is necessary to obtain correlation of air distance and ground distance In either case the aircraft instruments must be able to measure the atmospheric conditions and any necessary performance or stability and control parameters Document provided by SpaceAge Control Inc http spaceagecontrol com oo erkb shad nny bait al oat meee wir kitted ase Sool AEE MOST Se eee ae bin das aiin aiian ela eS lA Ba i LR P CWE CARTERET par ee wre tn eres are 8 4 1 Instrumented Range Operations A permanent range installation will provide a space position data acquisition and recording system Details concerning the equipment specifications should be obtained from the manufacturer or the rauge facility Ref 27 The range support instrumentation must provide means for conducting the test measurement and recording of all data data pro cessing unique to the range equipment and any necessary interface with the test aircraft The range timing system must be available to control the test sequence and provide corre lation of data recorded at various locations The range timing method may be different from the airborne time code generator and provision must be made to intograte the systems Various methods of synchronizing timing systems include a physical connection between ground station and the test aircraft using an R F link to impose time at one location on the ot
53. ed before conducting flight tests Ground vibration tests are conducted 4 to determine the blade natural frequency and mode shapes The blades and hubs all possible actual hardware are then placed in a whirl tower to confirm the ground tests Further vibration and stress data are then obtained from a restrained aircraft These tests provide information concerning stress distributions magnitudes of loads and boundaries for blade a compressibility or stall 4 Document provided by SpaceAge fetes NE VEP Oi EIEEEI tes PETERS ER VEET POIA Ea EAA Control Inc http spaceagecontrol com REIR oe niten cust tare Se RSA DI sic sh As a rule the preliminary tests do not flight conditions and it is necessary to obtain flight test data to accurately assess rotor performance stability and structural capabilities In some cases the instrumented rotor components from the ground test are available for the flight tests Such equipment will reduce the instrumen tation needed for the flight test and will produce the best comparative data for deter mination of effects of actual flight conditions New instrumentation must make maximum use of all test resulta to insure that the proper sensors are placed in the correct locations Occasionally a blade will be fully instrumented More common practice is to instrument the most critical locations for comparisons with design expectations and previous test results Most rotor syst
54. em Cycle counting is used to establish total accumulation Electronic circuitry is incorporated which calculates rate of accretion l during each cycle i The hot film anemometer is an electrically heated surface which is one leg of amp f wheatstone bridge network powered by the output of a high frequency high gain differential amplifier were bridge unbalance determines the amplifier output When a water droplet impinges on the sensor it is abruptly cooled The resistance of the sensor is highly temperature dependent and the cooling causes a bridge unbalance which is sensed by the 3 differential amplifier The amplifier applies sufficient power to the brilge network to return the sensor to equilibrium temperature The number of cycles indicates the droplet distribution and the applied voltage shows the droplet size Calibration data are then 4 applied to calculat droplet iaformation and liquid water content The frequency response 3 of the system is cr tical with respect to the distortion and attenuation of the droplet dats signal in the processing and recording portions of the system Large droplets or muitiple droplet strikes may cause data loss if the temperature does not recover before the next strike occurs Network noise must be minimized in order for the output freca small droplets to be recognizable _ Document provided by SpaceAge Control Inc http spaceagecontrol com not a AA ASR DTD REACH meet eee a pi tle e Peann DAA
55. ems have symmetrical parts so that it is only necessary to instrument typical components such as one blade one hub attachment or one control linkage When this is done consideration muat be given to any mass imbalance that may resuit agg Bee JENIN RS 5 1 Blades Se n ya The sensors placed on the blade must consider aerodynamics as well as structures The sensors must not create extra drag or reduce lift An aerodynamicist should provide guidance as to the best locations A stress analyst should be consulted to insure that the desired loads are being measured Equal consideration must be given to any wiring on the blade from the sensor to the recording system Significant aerodynamic effects can result from wires placed incorrectly on the lifting surface kE cones SOE i f a i Vibratory stresses are best measured with resistance gages The strain gages are bonded to the blade using the proper technique and the greatest possible care The gages are oriented to provide blade measurements of torsion in pitch flapping and in plane bending 5 1 1 Airflow In certain cases it is necessary to determine the nature of the airflow around the blade Visual displays such as tufts smoke or oil films provide qualitative informa tion and are most useful as a guide to the best location for the sensors The sensor location and data to be measured are provided by an aerodynamicist A common technique is
56. ent sources are most accurate when gensors are in exactly the same location on the structure 4 6 1 Sensor Location cw iiid ate on ert ak a ee pat en ee Le wets Dae wR Le ei Sensors should be placed on all components expected to be fatigue critical Special considerntion should be given to structures directly transmitting or receiving thrust or lift forces Examples would be tail boom mounting structures transmission mounts and wing or stabilizer attachments Specizic guidance on sensor location is not possible and the instrumentation must follow the directions of the stress analyst who will consider local stress concentrations operating environment and inter relations with other components or structure Riese nR 4 6 2 Sensors Load measurement is beat accomplished with a bonded strain gage bridge Particular attention must be given to the gage factor type of material being tested environmental temperature and conditions at the mounting location Bonding must be of the highest quality Calibrations can be calculated on the basis of sensor specifications and verified dynamically Pfs tetas tna Race arae EEE Document provided by SpaceAge Control neare Mave cache feredsis ature a Inc http spaceagecontrol co n p 4 7 Cockpit and Cabin Environment Tests may be conducted to measure the operating environment of the crew compart ments to insure that occupants can function adequately throughout
57. entation with data recording systems and ground support facilities are alse considered A summary of instrumentation requirements is provided along with recommended range accuracy and resolution A sample instrumentation management technique is also included 1 INTRODUCTION The definition of flight test will vary as widely as the activity of those who are pursuing the subject Perhaps the only consensus is that the vehicle be in free at mosphere as opposed to a wind tunnel or an enclosure Scale models unpiloted vehicles tethered or constrained vehicles are flight tested The type of power or even the absence of power is not a decisive factor Flight test may involve measurements rely upon opinions or simply be a demonstration of success or failure This document will deal with flight testing of helicopters where it is necessary to record data which describe the vehicle operation and response to specified conditions and maneuvers Individual sensors systems and recording devices must be incorporated into the test vehicles in a manner that will meet the data requirements within the known constraints The total instrumentation system is best designed by starting with the data requirements Accuracy data quantity reliability and physical characteristics of the test vehicle are first considerations Sensors can then be selected within cost availability and instal lation limitations Characteristics of the sensor are evaluated to det
58. eport No 0 0772 FR P 3 5 4 F E Jones Air Density and Helicopter Lift 1973 Joint Army Navy Aircraft Instrumen tation Research Report No 721201 P 16 24 5 D Belte F L Dominick J C O Conner US Army Aviation Engineering Flight Activity 1977 Helicopter Lift Margin System and Low Speed Performance Evaluation NUH 1M Helicopter USAAEFA Report No 73 01 P 14 and 78 6 Paul Spyers Duran Meteorology Research Inc Altadena CA Measuring the Size Concen aay and Structural Aspects of Hydrometers in Clouds with Impact and Replicator Devices P 3 7 W Kleuters dnd G Wolfer AGARD Advisory Report No 127 Some Recent Results on Icing Parameters P 1 1 through 1 10 8 A R Jones W Lewis Ames Aeronautical Laboratory Moffett Field CA NACA Research Memorandum A9C09 Apr 26 1949 A Review of Instruments Developed for the Measurement of the Meteorological Factors Conducive to Aircraft Icing P 2 11 ee Terie heer see ea te Sail ica SS aE Ts eed corte Ree EEI Cae Ast a l 9 R G Knollenberg The National Center for Atmospheric Research Boulder CO Journal fe of Applied Meteorology Volume 9 February 1970 The Optical Arran An Alternative to i Scattering or Extinction for Airborne Particle Size Determination P 86 90 i 10 R G Keller General Electric Company Aircraft Engine Group Cincinnati OH AGARD i The Propulsion and Energetics Panel 5137 as Specialists Meeting Icing Test
59. ermine recording system requirements Calibration schedules for sensors and systems can then be established E ee ee a eaa S A helicopter flight test installation requires considoration of some parameters N that are unique to the helicopter In other instances the instrumentation is common to other airborne vehicles however special attention must be given to items such as re cording range or sensitivity The instrumentation must provide data which will aliow the flight crew to establish flight conditions as well as record data for engineering appli cations 1 1 Test Vehicles The instrumentation discussed is directly applicable to helicopters In a more general sense it is intended for any vehicle which operates in a low speed omni directional flight regime The helicopters flight envelope introduces a need for special instrumentation For example it may be necessary to determine airspeed in all directions end the altimeter must be capable of operating over a large height band while retaining high accuracy for operatione near the ground Mechanical complexity is necessary to inte grate engines rotors and control systems Losses in the power transfer system are often small and are difficult to define accurately The mechanical problems associated with rotating masses introduces a need to measure many angular motions positions and torques The rotating n smbers cover the range from engine speeds to very low shaft
60. es by pitch links Aerodynamic and blade dynamic loads from the blades are also fed into tha pitch links as are static loads when the rotor is not turning Stress analysis will determine number and location of strain gages required The gages are mounted to measure vibratory tension and compression Strain gage specifications and mounting considerations have been previously Document provided by SpaceAge Control Inc http spaceagecontrol com Bie ase UA bits ii aida ia DIUM o a tai 4 1 aie ee Z F iS TN HE EE FS TE OES So DAE singgat we we Rea ee oe RR Roh ESR ER AEE RTS A VNR nee ree eS Teas e discussed In most designs the pitch link loads are transferred directly to the swash plate and can be measured at that location 5 4 Data Transfer The rotor data measured on the rotating parts must be transferred to the station ary data recording system in the aircraft Mechanical slip ring and brush devices are the most common method The strain gage signals are very low voltages and the slip ring assembly must not generate noise which will influence the data Factors which must be considered in the slip ring design are a Shaft speed and diameter b Slip ring material and surface condition hardness finish eccentricity c Brush material contact pressure and number of brushes The slip rings can be mounted
61. for the two unknowns Sota pom EI E oct See Sait FES S E Pe a5 a heb thio a a a Ss TERE e ap STM OS ee CET EEE RTE RE eR CURSE A ENRETE IS MU NY IRE Ce e FORWARD LEFT RIGHT REARWARD Figure 2 2 2 3 J TEC Airspeed Sensor At the inboard end of each tube near the hub is a vortex strut a wire of known diameter located just ahead of an ultrasonic transducer As air moves through the tube and across the strut a series f alternating vortices is created The frequency of these vortices is directly proportionai to true air velocity and is independent of density The vortices pass through an ultrasonic beam transmitter modulating it The modulation frequency is detected and is sent to its receiver where it is converted to an audio frequency signal The electronic processor and its case a box 12 7 om 5 in wide by 20 32 om 8 in high by 50 8 cm 20 in long weighs 3 18 kg 7 lb It converts the input audio frequency signals from the sensor to voltages and determines which two adjacent tubes have the greatest velocities The processor outputs two voltages proportional to longi tudinal and lateral true airspeed Typically the calibration is approximately 100 mv m s 50 mv kt Airspeeds are calculated within the processor The cross pointer indicator in the cockpit has a fixed display in the form of concentric circles 10 Kn 5 m s apart with zero located at the
62. formation concerning what the operational pilot will be seeing This standard instrument is often satisfactory for user or flying qualities evaluations 2 2 1 Swivel Head Test System To minimize effects of angles of attack and sideslip or pressure distortions near the fuselage a swiveling pitot static pzohe is mounted on a boom extending forward from the helicopter Ideally the boom length should place the sensor beyond the rotor downwash in hover However this is usually not practical When the boom cannot be placed on the nose other possibilities include wing fuselage or vertical stabilizer mountings Unusual airflow and vibration conditions should be expected in these locations Consider ation must be given to the natural frequency of the boom to prevent excessive rotor induced vibrations The boom is attached to the airframe on one end and supported by cables from the other end A typical boom is 2 5 to 3m 8 to 10 ft long and is constructed of 2024 T3 aluminum with a 4 45 cm 1 3 4 in outside diameter with a 32 cm 1 8 in wall tnickness The boom and pitot static probe combination can be precisely aligned with the aircraft axes by use of survey equipment The center line of the aircraft as determined from known airframe reference points is projected approximately 15 m 50 ft in front of the venicle The offset distance from the boom mounting point on the fuselage and the aircraft center lin
63. g aircraft reference system numbers FS Fuselage station longitudinal BL Buttline lateral WL Waterline vertical WI Weight of transducer and associated equipment in pounds to the nearest pound 4 CHANGE a AA DATE Effective Julian date of last change to each parameter Note Not required on i original list j i FLT First effective flight number for which the change is applicable 4 REMARKS Y REMARKS Reason for latest change or any other pertinent information relating to that particular parameter 1 4 4 4 4 i i j q 4 4 H Document provided by SpaceAge Control Inc http spaceagecontrol com is
64. geometric center and 50 Kn 25 m s at the outer ring The horizontal pointer moves up with increasing forward airspeed the vertical pointer moves in the direction of lateral aircraft motion The intersection of the two pointers indicates resultant vector airspeed d LOPAS 1000 The LORAS 1000 made by Pacer Systems Inc of Arlington Virginia consists of a sensor unit air data converter omni directional airspeed density altitude indicator and a contro panel The sensor consists of two venturi tubes mounted on opposite ends of a tubular rotor The venturis are connected to opposite sides of a differential pressure transducer A motor drives the rotor at a constant speed of 720 rpm in the horizontal plane to assure adequate dynamic pressure in the venturis independent of aircraft motion The air data converter combines the sensor unit outputs differential pressure and the corresponding angular position of the venturis with temperature and static pressure and outputs longitudinal lateral and resultant true airspeed Density altitude is also an optional output of the computer The system was designed to operate over an airspeed range of 25 m s 50 Kn true airspeed KTAS rearward to 100 m s 200 Kn forward and to 25 m s 50 Kn in lateral flight The system was also designed to be insensitive tc vertical motion and its method of operation is shown by Figure 2 2 2 4 RE ih ba BSS TRIE Rc tee ees Eh ane as ae seen ee 4 t i
65. haracteristics of the systems as they were tested may be obtained from references 16 through 21 TABLE 2 2 21 SUMMARY OF OMNI DIRECTLONAL AIRSPEED SYSTEM CAPABILITIES Longitudinal Lateral Vertical Angle of Airspeed Airspeed Airspeed Sideslip Kn Kn Ft Min t Deg O Rearward to 50 left to 40 rearward 40 left to 0 to 30 rearward to 40 left to 50 rearward to 50 left to 40 rearward to 50 left to 50 rearward to 50 left to 0 to 0 to 180 200 forward 50 right 5000 up 90 up Data shown are for the sensor mounted vertically Forward or lateral mounting will change the capability in the various axes With the exception of the Elliott rotor downwash will adversely affect performance Aeroflex Laboratories Inc Plainview Long Island New York was responsible for development of the true airspeed vector system TAVS The TAVS consists of an air stream direction sensor a true airspeed sensor a visual indicator and the associated electronics The corresponding sideslip angle in degrees and true airspeed in knots are available as DC signals suitable for recording on an oscillograph or magnetic tape system rhe airstream direction sensor consists of four hot wire senso mounted on top of the airspeed stream tupe The bolometers form an e triacs of the airflow direction with respect to the longitudinal axis of the stream tuba Airflow at an angle to the turbine duct causas the right and left bolometers to be cooled unequall
66. hath te Stes a adrian bo ae tine eae ae Seg Sern age hae 2 2 Airspeed and Relative Wind Direction Standard airspeed systems are oriented for conventional level flight These systems sre fixed pitot static differential pressure probes They are usually designed and arranged to show a minimum error at cruise airspeed The threshold is relatively high and the systems are often unuseable at indicated airspeeds below 15 m s 30 Kn In level tlight sideslip effects on position error normally increase with airspeed Sideslip angle up to 5 degrees will not usually introduce noticeable errors This is an important consideration since most helicopters have sideforce cues which allow the pilot to stay within these limits Very steep climbs or descents will cause noticeable shifts in the position error for angle of attack or sideslip and separate calibrations are accomplished for those flight regimes Refs 13 and 14 Standard cockpit indicators normally have 2 5 m s 5 Kn increments and the instruments are not calibrated but are accepted with a specification accuracy tolerance Thus the accuracy of a particular instrument in a given aircraft is not known When the standard sensors are used for test data it is common practice to mount a test indicator in the pilot s instrument panel The test indicator is calibrated and can be read to 0 5 m s 1 Kn increments During the test data will also be taken from the copilot s standard instrument to obtain in
67. he data are to be used for power correc tions due to accelerations in the various axes the sensors are placed as near the center of gravity as possible to minimize the effects of angular motion During systems testing or handling qualities evaluations it may be necessary to measure the total acceleration linear plus angular at a component or at the pilot s station In these specialized cases it may not be required that all axes be instrumented The omni directional flight capability of the helicopter imposes requirements somewhat different than fixed wing aircrast Sideward and forward acceleration capability are nearly equal and may be up to 1 G Normal acceleration seldom exceeds 3 G The accelerometer data should be very accurate since small errors can introduce large variations in the performance calculations In Pee m DSB een IZA Uh AEL ANR D AEGI A RPE TER ho eon Berii o Meee EER oss ee addition to high accuracy the sensitivity and frequency response must be sufficient to i A E record rapidly changing conditions during maneuvers A large variety of accelerometers are x available for measuring this range of acceleration The requirement to measure accelera 4 H tion in a frequency band which includes static accelerations can be satisfied by the use a a of 3train gage or piezoresistive accelerometers Although both types provide a lower A a amp frequency response of D C the upper limit van vary from 600 to 8000
68. her recording location and recording a common standard source at each location Communications for test control are usually voice transmissions over radio or telephonic instruments Precise data information is communicated over wire lines or radio or tele metry links The atmospheric measurements at the range will usually include information necessary to correct the raw data obtained from the acquisition system Parameters such as noise humidity pressure and temperature are necessary for assessment of aircraft performance and are usually measured by aircraft sensors These data may be telemetered to the ground station and recorded with the space position data Range facilities have various data processing capabilities Typical equipment provides data readout transla tion format conversion and automatic display Commonly used space position systems are optical radar doppler and laser A limited understanding of how the different systems operate is necessary to assess the instrumentation requirements Ballistic plate cameras record aircraft images on a glass emulsion plate These cameras may be fixed or may be tracking devices such as the Fair child Flight Analyzer Askania cinetheodolites track the aircraft and make a film record of the azimuth and elevation relative to a known set of coordinates Ribbon film cameras such 6 the Bowen Knapp track the aircraft and record the image along with fiducial markers which are projected onto the film Re
69. here may be a requirement to compare the system input to the computer with measurements from an independent test instrumentation system lypical comparative parameters include airspeed angle of attack aircraft attitude or acceleration A completely instru mented test aircraft as previously discussed will provide adequate comparative data In other cases the data requirements must be carefully studied and instrumentation added as required Generally measuring the fire control system is required to insure that the proper functions are being recorded i 7 5 E rnal Noise Noise from weapon systems must be considered from crew exposure and from aircraft detection by ground personnel Crew stations are instrumented to measure noise levels Ground recording stations are established in a grid around the helicopter to record the noise for different flight and firing conditions During forward flight the helicopter is flown over a prescribed course through an instrumented range Additional noise measuremont discussion is presented in section 8 2 eee te Sr aS peed meee E i e E a a DS cee cate 7 T SA EZE ESAR 3 cane AUA S rae tele aks dete Fela Gas Contamination i Toxic or explosive gas from the weapon way enter the crew compartments and engine inlet or may impinge upon the airframe or rotors Measurement of crew compartment contami nation was discussed in the airframe section Significant amounts of gas in any
70. ines distance by phase comparison of the transmitted and return signal With a doppler system a signal is radiated by a transmitter on the ground This signal is received on board the aircraft and retransmitted at a different frequency by a transponder At least three receiver stations on the ground are needed to receive the reference frequency and the retransmitted frequency These two frequencies are electron ically subtracted and the difference is the doppler frequency at the particular receiver The doppler frequencies are used to calculate the three dimensional position in space Additional receivers allow statistical techniques to be used in accuracy analysis The most recent development in space position equipment utilizes a laser tracking system The system radiates a short wave length signal which ic highly collimated and power is adjusted as a function of range The system operation is much like the tracking pulse radar system The accuracy is extremely good with capability for high density data that can be transmitted in real time or recorded on magnetic tape Tracking error is minimized by use of a retroreflactor element installed on the aircraft which provides a strong return signal that is tracked automatically In addition to improved tracking the apparent range is independent of the many atmospheric variables which must be corrected for in radar systems The Sylvania Electronics System has developed a self contained unit which is va
71. ing climbs to service ceiling Radar altimeters are often required for operations less than 300 m 1000 ft above ground level Test system static and pitot sources are placed in a location which will minimize effects from aircraft and best reflect the true atmospheric conditions When possible these sensors are placed on a nose boom The static pressure is connected to both cockpit indicators and instrumentation transducers For most altitude applications a standard indicator is used with 6 m 20 ft resolution This type of indicator is generally accept able for pilot information The instrumentation system altimeter can be a strain gage pressure transducer capacitive transducer or other suitable transducer The capacitive transducers come in both analog and digital output formats Temperature effects can be sizeable aud therefore should be quantified for necessary correction by appropriate cire t 5 enclosure of the transducer in a temperature controlled oven or data manipula tio ng analysis Care should be taken to provide any necessary preflight warmup time i esu transducers From one to thirty minutes may be required for proper stability v spoper installation and appropriate data correction accuracies of better than 134 Pa Peta tn of mercury can be realized which meets the requirements of most applications Very accurate height above ground information is often needed during hover and toef ud landing tests and a
72. ing for Aircraft Engines London Engiand Apr 3 4 1978 Measurement and Control of Simulated Environmental Icing Conditions in an Outdoor Free Jet Engine Ground Test Facility P 7 2 through 7 4 11 J D Hunt SVERDUP ARO INC AEDC Div Arnold Air Force Stn TN AGARD The Propul sion and Energetics Panel 5137 A Specialists Meeting Icing Testing for Aircraft Engines London England Apr 3 4 78 Engine Icing Measurement Capabilities at the AEDC P 6 3 through 6 10 eee ie pet 12 USAAMRDL TR 75 34A Volume 1 Design Criteria and Technology Considerations Develop ment of an Advanced Anti Icing Deicing Capability for US Army Helicopters Eustis Direc torate US Army Air Mobility Research and Development Laboratory FT Eustis VA P 134 142 13 K R Ferrell Cpt W J Hodgson Air Force Flight Test Center 1964 YCH 47A Category I Performance Stability and Control Tests Report No FTC TDR f 3 36 P 19 and 89 14 K R Ferrell J Shapley Jr J Mishlof US Army Aviation Systems Test Activity 1970 Wind Tunnel and Flight Evaluation Rosemount Shielded Pitot Static Tube Model 850N USAASTA Report No 68 12 P 11 13 15 K R Ferrell B Boirun Cpt G Hill US Army Aviation Engineering Flight Activity 1977 Low Airspeed Sensor Location Tests AH 1G Helicopter Final Report USAAEFA Report 75 19 1 P 8 12 16 K R Ferrell A Winn J Kishi B Jefferis US Army Aviation Systems Test Activity 1973 Flight Evalua
73. inlet Cockpit instruments are usually provided for in flight recording of pressure and temperature and should there be multiple seneors a switch should be provided to allow the flight crew to monitor the data 3 4 1 Inlet Pressure Most engines are delivered with at least one total inlet pressure sensor in stalled For a well defined engine or for a cursory performance evaluation this may provide sufficient information A single sensor is not satisfactory for rigorous perfor mance tests or for dealing with a new engine or installation To obtain data which wili show distortions and provide construction of pressure profiles it is necessary to use several sensors mounted on a rake and placed in a suitable location in the inlet The reke will best show engine inlet conditions when it is placed near the compressor face The number of sensors on the rake will depend on the data requirements physical nature of the inlet and the recording system capability Accuracy of the profile and distortion infor mation is very sensitive to the number of probes and the probe array Ref 22 Struts or any other physical characteristics of the inlet will influence the flow and may change engine performance A typical inlet rake is shown in Figure 3 4 1 1 tener AOR TEATS SIT IMMERSION RAKE TOTAL staric Figure 3 4 1 1 Inlet Rake Both static and total pressure sensors are required and should be located at the same engine
74. ion and data reduction software WORD NO Software calibration channel assigned by F T E or programmer in coordination with instrumentation engineer PCM FM CHAN PCM main frame location or FM band center frequency D suffix indicates gital channel Decimal notation indicates bits used in split digital channels Assigned by instrumentation engineer in coordination with flight test engineer CALIBRATION POSITIVE POLARITY Direction or sense of physical input that corresponds to increasing counts or frequency Engineering value at ZERO COUNT Engineering unit value that corresponds to zero 000 PCM counts or lower FM band edge PER COUNT Slope of relation between engineering units and PCM counts or FM subcarrier gain Note This also corresponds to the resolution of the parameter For non linear calibrations calibration function coefficients can be substituted MAX COUNT Engineering unit value that corresponds to PCM full scale or FM upper band edge PCM counts at ZERO VAL PCM count or subcarrier frequency that corresponds to an engineering value of zero if applicahle i e zero roll rate MAX VAL PCM count cr subcarrier frequency that corresponds to a maximum positive or other specified engineering unit value i e full 100 percent control deflection Note This does not necessarily correspond to the engineering value at Max Counts R CAL PCM count or subcarrier frequency that results from activation of a syste
75. ion deck listings 5 A brief chronological log will be maintained of all changes their effective date and flight and the reason for the change Detailed instructions for completing the form This contains detailed instruc tions for completing each entry on the form If a particular entry is not required for a parameter enter NA not applicable in that block For some parameters it may be desirable to substitute information other than that described here In that case inform the instrumen tation engineer so that future instructions can be updated Typing Information This form is arranged for standard elite type spacing 12 characters per inch 6 lines per inch with 14 space vertical spacing except for heading information The maximum number of characters per column is listed by each block title in the following instruc tions The form size is 11 x 17 inches and can be typed in standard size typewriters by folding it on the line between the major headings of CALIBRATION and SIGNAL CONDITIONER The entire form does not need to be retyped each time corrections or changes are made Correction tape can be applied over previous entries The only requirement is that clear reduced size copies can be made NOTE Great care should be taken in typing and proofreading this form as it is used by several people for a variety of purposes A single mistyped or misplaced character can have a significant impact Instructions for completing each
76. ion to overall noise production Noise measurements for flight regimes other than hover may be necessary to eval uate speed effects on noise propagation The most important effect is the impulse noise generated by biade tip vortex interactions The most common method is to fly over or near the hovering ground matrix of microphones An aiternate in flight technique has been developed ref 25 With this technique microphones are placed on a pacer aircraft The impulse noise from the rotor may be directionally sensitive and a lateral displacement of the microphones may be advisable The pacer aircraft is fitted with an automatic recording device or equipment that will transmit the signals to a ground station Provision should be made to adjust the instrumentation gain to optimize the signal to noise ratio The typical peak pressure will vary from 10 to 500 Pa 1 45 x 10 to 07 PSI with the maximum occurring at high advancing tip mach numbers iu forward flight The noise of the pacer aircraft ia obtained prior to the test and taken into account either through instrument adjustments or later in the data analysis 8 3 Thrust Hover performance requires measurement of the thrust that the helicopter gener ates for a given power setting This is usually accomplished by restraining the aircraft to the ground and measuring the various forces generated as a function of power or varia tion of thrust devices To be effective the thrust stand must have the cap
77. istics around the helicopter are extremely i complex for other than high speed flight and it is difficult to find a sersor location a which is free from aircraft disturbance or contamination While it is expected that ice will accrete on all parts of the aircraft it is not practical to measure ice thickness on blades or other rotating parts It is common practice to paint or tape the blades in a r grid which identifies span and chord locations Photographs are then taken to establish x patterns and amount of ice accreted Those determinations are correlated to the atmos pheric conditions and accretion measured on the fuselage or ovher non rotating parts Droplet size can be determined by various types of impact measuring devices Slides coated with oil gelatin or carbon are exposed to the airstream for a short period of time The droplets are either captured by the surface or leave marks representative of their size Examination under a microscope allows determination of size and distribution Another technique involves a water sensitive tape or paper which is continuously moving behind a slot exposed to the cloud Ref 6 This provides a time history of the droplets being encountered Droplet size can also be determined by the rotating cylinder method This method exposes cylinders of various diameters to the airstream with their axis perpendicular to the airflow The cylinders are rotated siowly so that the ice build up is uniform Th
78. leed air can often be obtained from engine test cell data with bleed air on and off Cockpit or cabin environmental control systems are often combined electrical airflow devices Cooling may be obtained by circulating outsi ia or cabin air through a refrigeration unit This unit may be electrically driven or may use engine bleed air Previously discussed instruments can be used to obtain power and airflow data 3 9 Power Plant Controls Power plant controls include those in the cockpit and at the engine Reciprocating engines usually have cockpit controls whose positions need to be measured while turbine controls are normelly on off or three position devices Automatic engine controls are often instrumented to provide correlation data and insight to engine performance char acteristics Engine and cockpit control relationships are needed to evaluate dynamic rasponse and engine rotor capabilities Potentiometers or microswitches are generally used as sensors for these applications The amount of motion can be expressed in degrees or linear measurement around the arc for a control which moves about a fulcrum Calibration of control motion is generally done by using an inclinometer to measure the angular motion in degrees and then by measur ing the radius of the control arm Attention must be given to properly identifying the radius of the oe arm The distance from the center of hand contact to the fulcrum is most often used 3 9 1 Cockpit Contr
79. lifier and by using low noise cable The output of a charge amplifier is strictly a function of the sensor charge output the amplifier feedback capacitor and charge noise generated by the cable By placing a low gain miniature charge amplifier near the sensor and then amplifying the resultant voltage output with the instrumentation signal conditioning low noise measurements can be made This low gain configuration helps in suppressing the triboelectric noise cable charge noise and eliminates cable capacitance effects A number of manufacturers produce suitable sharge amplifiers 4 6 Loads et ae E E hela Seg oe eee a A IERE TA hae Tie i Sata E The structural loads demonstration is conducted in conjunction with the early vibration testing During the performance or stability and control testing the limits of the envelope will be reached and new conditions or maneuvers may be attained Loads data can be used to evaluate the hardware suitability under mission operating conditions allow comparison with derign information and contribute to fatigue lite calculations Loads instrumentation is extremely critical with respect to sensor location and number of sensors Analysis of design information bench or component testing resvlts and previous flight tests will suggest critical locations Models can also be constructed of materials which will visually show stress concentrations Comparisons of data from these differ
80. luation of the methods now available to measure these rotational velocities centers around the magnitude and transitory nature of the velocity Those of high or low speed with little short term variation are easily measured but rapid changes in velocity must be given special attention In general less transient parameters are handled by measuring the frequency of rotation in a rather direct fashion As an example a constant rotor speed is often measured by outputting the rotor tachometer generator to a frequency to D C voltage converter This provides a D C voltage level proportional to rotor speed While this technique provides good results with little or no transient rotor conditions large rotor speed variations can result in sizeable measure ment errors By providing high resolution high sample rate period measurements of variable low speed shaft parameters problema of response to rapid frequency changes and or invalid data averaging associated with frequency to D C measurement techniques can be eliminated Accuracies greater than ti are possible Consideration should be given to measurement repetition rate master clock frequency etc required for the particular transient condi tions present 3 1 1 Engine Speed Engine speeds usually vary from high compressor or turbine speeds to lower shaft speeds following gear reductions Measuring internal engine speeds with a test system is difficult at best and may not be possible in some cases
81. m Usually it is necessary to measure pitch roll and yaw attitudes Sensor location relative to the aircraft body axes must be precisely known For tests near the ground photographs showing flight path and an earth reference can be used to measure the attitude Most teste sre sufficiently above the ground that ground mounted cameras cannot be used Photogrisiis from chase aircraft seldom show true angles and thus cannot be used for engineering data With the proper equipment celestial bodies in conjunction with earth position can be used to determine the aircraft attitude These photographic and optical techniques require special equipment The data are difficult to process and cannot be used in certain atmospheric conditions For steady state condition pendulum type sensors cz use earth gravity to measure the relative position of the aircraft Acceleration effects render such a system unuseable for most flight test purposes Gyroscopes mounted on each of the aircraft axes provide the best approach to obtaining the aircraft attitude Certain helicopter character istics alter the requirements somewhat from those needed in a fixed wing application While some helicopters are fully aerobatic pitch attitude will seldom be more than 60 and roll is usually within 100 In several tests the helicopter is required to yaw 360 at hover and low speeds The gyroscope data is usually recorded continuously for times of less than one minute which should be
82. m Fire Control Wind Sensor Report Final Report USAAEFA Report No 75 13 2 P 10 14 23 F Stoll J W Tremback H H Arnaiz 1979 Effect of Number of Probes and their Orientation on the Calculation of Several Compressor Face Distortion Descriptions NASA TM 72859 P 7 9 24 K R Ferrell Maj W Welter 1967 US Army Test Office Engineering Flight Research Evaluation of the XV 5A Lift Fan Aircraft Pt II Performance Finel Report USATO Report No 62 72 2 P 72 75 25 F H Schmitz V Duffy 1977 In Flight Measurement of Aircraft Acoustic Signals Advances in Test Measurement Volume 14 Proceedings of the 23rd International Instru mentation Symposium Las Vegas Nev 26 Capt G D Tebben USAF R K Ransone 1965 Evaluation and Checkout of the Air Force Flight Center VTOL Test Stand Feb 1965 AFFTC TR 34 37 2 8 27 R G Culpepper R D Murphy E A Gillespie A G Lane Aug 1979 A Unique Facility for V STOL Aircraft Hover Testing NASA TP 1473 P 7 28 28 Trisponder 202A including 202 RO6C Del Norte Technology Inc b Apr 76 29 F D Schick An Electronic Method for Measuring TakeOff and Landing Distances Society of Flight Test Engineer Symposium Proceedings 4 6 August 1976 30 W Y Abbott Del Norte Space Positioning System Development Report and User s Manual Sep 1976 USAAEFA Technical Note 77 64 P1 4 31 W Beech et al Air Force Flight Test Center Propulsion System and Performance Evaluation of
83. m self test feature An acceptance tolerance can also be specified for preflight purposes DATE Latest parameter calibration Julian date last digit of year day of year i e 3062 21 February 1975 SIGNAL CONDITIONER NO Index number assigned to signal conditioning equipment used for parameter CHAN Subdivision number in multiple channel signal conditioning unit FILTER Basic filter characteristic and corner of pre sample filters used in analog signal conditioning i e BLP 10 Butterworth low pass filter with 10 hertz cutoff TRANSDUCER NO Arbitrary index of transducers TYPE MODEL Manufacturers name plate type model or part number or transducer SERIAL NO Manufacturers name plate serial number of transducer SIGNAL Code identifying the form of the signal generated by the transducer The code is as follows Apalog Signals A nalog High Level Bipolar AHLP DC Analog High Level Unipolar Positive AHLN DC Analog High Level Unipolar Negative ALL DC Analog Low Level Bipolar ALLP DC Analog Low Level Unipolar Positive ALLN DC Analog Low Level Unipolar Negative AC Signals Control Transmitter CT Control Transformer CR Control Receiver Y1 X1 Synchro Transmitter Receiver Line Polarity Y2 Yl Synchro Transmitter Receiver Line Polarity Y3 Z1 Synchro Transmitter Receiver Line Polarity TG2 Tachometer generator number of poles Document provided by SpaceAge Control Inc http sp
84. m time Centralized location of the necessary equipment reduces time and eases checknut or correction procedures Maximum accessibility is gained by placing racks away from the sides of the compartments and by using a minimum of closed panels Routing of cables snould consider electro magnetic interference as well as allow visual and electrical inspection The racks containing the instrumentation and instrument mountings must be designed to withstand specified loadings The instrumentation should be able to withstand forces greater than the occupant seats or restraints t insure safety during an accident Typical design is for impact forces of 20 g s in each axis Wires cables ur other restraining devices should not present hazards during normal operations around or with the equipment The weight and location of each piece of equipment must be known The instrumen tation engineer should coordinate with the flight test engineer to consider the total weight of the instrumentation with respect to performance capability of the aircraft and location of instruments or components with respect to the center of gravity and inertia Weight and locations are often critical for small test vebicles Common practice is to write the weight on the larger pieces of equipmeat This provides a rough accounting during the installation When the installation is complste an aircraft weight and balance ia required to account for wiring and small miscellaneous items 2
85. ment provided by SpaceAge Control Inc http spaceagecontrol com aramea k e EVANTA HAN AB ta h ra tA i el abel f 4 y a REET SE Te eS eng TM perce Tier iw Do SEP g Ste s scons este as The noise measurements are usually taken at different azimuths at specified distances from the helicopter while it is on the ground or hovering Piezoelectric type microphones are commonly used An amplifier system is interfaced to the microphone with preamplifiers used in some cases The circuitry can be designed to meet program specific frequency response characteristics and provide the desired output by incorporating the necessary filters The amplifier output data is recorded by magnetic tape or oscillographs for later analysis Whether this is recorded in a direct or FM format on tape care should be taken to insure that the frequency response is not degraded by the magnetic tape recorder With oscillographs the galvanometers must be selected and set up to insure that desired signal information is not degraded Quick look capability can be provided in a fixed base system by use of a real time narrow band spectrum analyzer in parallel with the recorder With the spectrum analysis the number of samples averaged and therefore the degrees of freedom must be carefully chosen to obtain a desired confidence limit This scheme allows the driving inputs at different frequencies to be quickly assessed for contribut
86. n mounted and can be transported to road accessible sites This system incor porates a mini computer and assorted dais handling equipment 8 4 2 Remote Site Operations The versatility of the helicopter allows testing in remote sites where ground support and perhaps sven a runway do not exist All test equipment must be portable use a minimum of power and be able to operate in an adverse environment Systems which are crude in comparison with range instrumentation can produce surprisingly good data when used with care In most cases some of the previously described optical systems are used because of simplicity and low power requirements When optical systems are used to record data visual theodolites can produce quick look data and in extreme situations may be i i 5 i j a j i d i d Document provided by SpaceAge Control Inc http spaceagecontrol com 3 ash etl ilhi eee Bateman te Ah irua ste ard s ai TA BUAR E ni atadas athuutisredels i eS ie as Aerie tence ar cues pS oi the source for final data The greatest flexibility in remote site operations are achieved with airborne acquisition sys ems The simplest system is a camera mounted on the aircraft which records torrain or markers during the test maneuver The markers are usually runway lights or distances along the flight path These markers must be carefully surveyed a
87. n the power required terms During a propulsion system analysis these powers affect the total engine power available and in the case of airflow taken from a particular engine statio may influence the thermo dynamics of engine operation The systems are often complex and redundant and each must be carefully analyzed to determine sensor location so that the proper power is being measured Electrical power from the alternator is determized by measuring power directly or indirectly from measured values of voltage current and the phase angle between voltage and current Standard engine sensors normally are provided and the signal can usually be recorded by the test data system High accuracy requirements for power usage by any com ponent will dictate installation of test sensors Hydraplic system power may require measurement of power to drive the pumps or the flow of hydraulic fluid The pump may be driven by engine gear box rotor transmission electrical power or bleed air Gear trains or drive shafts require a speed and torque Measurement as previously discussed Provided the density of the hydraulic fluid is known the hydraulic mass flow may be determined from the measurement of volume using a method similar to that for determination of fuel mass flow as discussed in paragraph 3 7 1 Electrical pumps need voltage and current measurements Bleed air flow is measured by pressure and temperature sensors Another determination of power loss to b
88. nd the data accuracy is highly dependent upon interpretation of the film Correlation of the film with other recorded data is also very difficult An aircraft portable radio ranging system has been developed by Del Norte Inc Euless Texas Ref 28 29 and 30 The system is solid state compact and has a range of 4 8 km 3 miles This system has a distance measuring unit UMU which controls all operations a master unit MU which transmits and receives all signals and remote units RU which receive signals from the master unit and retranemits a signal The airborne equipment operates on a 24 volt D C power while the ground units are powered with 12 volt D C battery power The system is line of sight and operates on radio frequency signals The DMU and MU are located in the aircraft and the remote units are placed along the test area As many as eight ground units may be used The remote stations require no ground support personnel and for simple runway distance a single remote unit can be used without a site survey The time for the signals to travel to the remote units and return is measured and provides slant range distance Data from individual ground stations can be examined for random points dropouts or multi path interference Stations which provide best data quality can be used to calculate horizontal and lateral displacement to within tl m 3 ft However at relatively low heights above the ground small errors in range create extremely
89. nge on airframe structure stabilizers and tail rotors Those exhausts may cause surface heating which can be measured with thermocouples In addition ingestion of gun gas products or missile propellant products can have a j aevere impact on engine operation The hot gases can cause airflow disturbances which 4 affect atabilizer lift or tail rotor thrust and cause a change in stability and control eerourence studies will determine the necessity of uging anemometers to measure the exhaust 7 3 Ejected Material Gun systems usually eject shell casings links or cartridges into the free stream Still air patterns from ground tests and expected airflow are used to estimate the in flight dispersion patterns Cameras are mounted to photograph the ejected material The camera installation must not disturb the airflow or change flutter or structural character istics of the menpe gt o which they are attached Camera speeds of 400 frames per second will provide data suitable for a pattern analysis he mee ge vp te Ler POLEC arene me SE rss 7 4 Fire Contro Systems lt Complex fire control systems utilize data inputs concerning the atmosphere target and aircraft conditions These inputs may be from standard aircraft sensors or they may be an integral part of the weapon system The inputs are fed to a computer which aims the weapon or makes corrections prior to firing or during the missile flight In either case t
90. ol Inc http spaceagecontrol com ERETO POTEO atin pill RN AAG ee th coe ci Ge Uy Cn Ph DA SI T Pe REL ATTA een IST Npe SO Her Ie aa erent eT e e we NT PR AL ot Epox FP A ARN a A gi 3 4 2 Inlet Temper ture As in the case for pressure an inlet temperature sengor is often standard with the engine and may be used as a suitable data source in some cases During hover or in low speed omni directional flight the engine exhaust may be trapped in the down wash and be re circulated into the engine Heat from transmission systems can also find its way to the engine inlet These heat sources can cause large inlet temperature rises with dramatic effects on engine or aircraft performance Ref 23 For re ingested gas the temperature may be uniform across the inlet while for radiated heat it may be concentrated in a parti cular sector of the inlet Complete inlet temperature data require a rake with probes spaced at different levels and azimuths Temperature sensors can also be placed on the pressure rake as previously discussed The number of sensors is established by the degree to which the profile must be determined Hot gas re ingestion may cause temperature rises of 50 C and the flow is very turbulent which causes large rapid fluctuations and dictates a high response characteristic for the sensor A suitable inlet temperature sensor is a chromal constantan thermocoupl
91. ols Power levers or twist grip throttles are instrumented with position transducers to show position from closed to full open For on off controls microswitches are used to record the positions Microswitches or position transducers can also te used to show activation of engine trim devices Strain gages are placed on the controls to measure forces applied by the pilot 3 9 2 Engine Controls To determine lost motion and delay in activation the controls at the engine are instrumented and results are coupcred to control motions in the cockpit Fuel control levers can be measured in terms of angular or linear displacement A very sensitive sensor Document provided by SpaceAge Control Inc http spaceagecontrol com PEE EE tokiai o rimes foe an vae Tiete ea Uone SRST IE HA ALL er AAAS TAME Ae a iea baasai AAT Aata aLi a T ES e ARD EU Ey te na sities Rode AA ek sal ae Le Lied a eee eee net el tented L E must aa ERT EERTE EN ORTAS be used since the displacements may be very small Engine controls such as droop compensation are not controlled directly by the pilot and motion must be correlated with items other than cockpit engine controls Engine governor inputs to these can be instru mented as previously discussed 3 10 Engine Vibration It may be necessary to measure either vibration generated by the engine or the vibration from the airframe to the engine Transducers are place
92. otions which are generated by 2 54 cm 1 in control inputs that are held for are one second Typically a rate gyro with 30 sec in pitch 100 sec in roll and t60 sec ie a in yaw is used the format cf the electrical output of the rate gyro motion will vary but i p gt a high level single ended output is most often used 4 3 Angular Acceleration a Angular acceleration data is needed to assess the helicopter controllability and for use in aircraft energy calculations Angular accelerations are often computed by differentiating the rate gyro output When necessary accelerometers are usually mouuted at the aircraft center of gravity and are aligned with the pitch roll and yaw axes The alignment and location of each sensor must be accurately established and recorded The helicopter angular acceleration can be up to 200 sac Stability and control systems input must be considered and the sensor is usually sized for systems off which produces the highest acceleration The pitch roll and yaw accelerations are inversely propor tional to aire aft moments of inertia and therefore are usually highest in the roll axis Presently available angular accelerometers are highly susceptible to vibration contami nation making mounting critical i 4 4 Linear Acceleration a Linear accelerations are required for energy analysis in performance testing and for certain stability and control tests When t
93. otor ground effect is usually quite strong which mandates that heights within one rotor diameter above the ground be deter mined very accurately Considerable noise is generated by the engines power transfer mechanisms and other rotating parts In addition the rotors contribute significant aerodynamic noise that may include a wide range of frequencies and magnitudes Document provided by SpaceAge Control Inc http spaceagecontrol com Types of Tests The type of test being conducted will significantly influence the instrumentation requirements A typical listing of tests is shown in Table 1 2 1 i 2 TABLE 1 2 1 Typical Helicopter Flight Tests Performance Hover Performance Take off Performance Climb Performance Vertical Forward Flight Level Flight Performance Maneuvering Performance Acceleration and Deceleration Turning h Dive Recove y Return to Target Terrain Following X Autorotational Descent Performance ane Landing Performance Handling Qualities Control System Characteristics Control Positions in Trimmed Forward Flight Static Longitudiral Stability Static Lateral Directional Stability Maneuvering Stability Dynamic Stability Controllability i Ground or Deck Handling Characteristics 9 Takeoff and Lending Characteristics i w Slope Lunding Characteristics A w High and Low Speed Flight Characteristics y rei N Power Management E i Mission Maneuvering Characteristics F t Fftects of We
94. ow Since the positioning torque is low the sensor must have low friction bearings and extremely good balance Potentiometers and bending beam strain gages are the most common methods of sensing angle of attack and sideslip vane positions Document provided by SpaceAge Control Inc http spaceagecontrol com NELERTE ES res oot AN gt aAA i t 4 t t i i i A t ee e e e e ved iarten story an h Ti ANID EAAS aE BRL ta RTS NPIS CE CRN Ht RENTER ITY SE TT HN PETE ET CO TEE A Pitot static systems are usually volume balanced to eliminate airspeed indication errors caused by differential pressure lags in the two circuits during climb or descent The pitot and static circuits are trial and error balanced by applying a pressure or vacuum to both the pitot and static sources simultaneously The pressure or vacuum is then bled to ambient pressure at a constant rate 120 m s 4000 ft min and the pressure differ ential read on the installed sensitive airspeed indicator Care should be taken to avoid over pressuring the airspeed indicator particularly in the negative direction A known volume 160 cc 10 cu in is then added usually to the pitot side and the process repeated The final volume to be added can then be calculated by linearly extrapolating interpolating the change in differential pressure caused by the known change in volume Usually two iterations are sufficient to balance the s
95. re engine output shaft speed is a irequeacy to D C converter The output of the converter is a voltage proportional to the speed and is recorded by the instrumentation system 3 1 2 Drive Shaft Speed The drive shaft speeds which must be measured will depend on the test require ments and the physical nature of the test vehicle In some instances it may be necessary to know the speed while in other instances the power being transmitted is of prime importance An example of the first case is determination of rotor speed by measurement of input shaft speed to a transmission The second case arises when power must be known for each component in the drive system Transmission losses can only be established by measuring input and output power This necesaitates a shaft speed measurement A magnetic sensor and chic a system similar to that used for the engine output shaft is the most common method 3 2 Engine Torque Engines commoniy have a torquemeter which can be incorporated into the test instrumentation system the wide variety of aircraft types requires that the instrumen tation system have great flexibility for interfacing with engine torque sensors Rather than measuring torque directly it is more common to sense some characteristic which is proportional to torque The sensing devices in use include monitoring electrical per meability of the shaft optical measurement of the shaft twist and strain gages for torsional measurements Appropria
96. re an instrumentation system with a wide range of dynamic response Design emphasis in this area can have a most significant impar on the system The parameters can b divided into low frequency atmospheric conditions and aircraft operation medium frequency aircraft motio and response and high frequency vibration and structural loads mesuurements The data parameters stould be grouped by dynamic response and maximum use chould be made of electrical filters and multiplexing Decisions must be made regarding the form of the data recording anc the data processing methods to be used Ref 1 The minimum requirement will be dictated by the tests The most common method is recording elgctrical signals on magnetic tape The tape may be on board the aircraft or the data may be transmitted to a ground station It is Document provided by SpaceAge Control Inc Dada at co na on tae ss E ERENER e CASS ATECA A OEREN sisi http spaceagecontrol com en SAFO RTE ARES CL YO A TS TR NT MEE yee om T a a ne i ea Ne NL SEEN LE SA ET PEEN E E E ee i A ER L T E E E OFE E E i ey EE E SS Pee Te ae eg Be a ae a erat Ser 6 oT Paes tO oF SOE OR a SEE Set Ir he i not unusual for both methods to be used Documentation of the data can be accomplished by use of a voice track in the recording system or with written notations by the instrumenta tion operator Pro
97. rovides procedures and instructions for preparation and use of an instrumentation form during formulation and conduct of an engineering flight test program Emphasis must be placed on the bookkeeping to insure that the instrumentation configura tion and status is correct for any proposed test This information generally typifies the approach of the flight test community however it should be modified to accommodate specific procedures instructions that may vary widely These instructions are for a pulse code modulation PCM data system which is in most common use today The purpose of the form is to consolidate and standardize all of the airborne recorded instrumentation project information to eliminate common coordination errors The sample form shown and these instructions should be modified to meet individual requirements A form should be com pleted for each project using airborne recorded instrumentation The instrumentation or data systems office is the proponent for the form and is responsible for maintaining the status and instructions current Chronologically the following actions are taken by the following responsible individuals to prepare and maintain the Recorded Instrument Parameter List s Flight Test Engineer F T E The F T E requests airborne instrumentation from the instrumentation or data systems office The request will contain schedules controls displays instruments and otker information The majority of information
98. rsions or shears More than one tower allows comparison of the air upstream and downstream from the helicopter 4 EES i 8 1 1 Wind Speed Accurate measurement of small rapid wind speed changes or local air velocities are best accomplished with a hot film anemometer An ultrasonic sensor could also be adapted to this application Lower response or time averaged data is usually obtained from a vane mounted pressure transducer or a cup anemometer Rotating cup anemometers carefuly constructed to minimize weight and friction can be sensitive to speeds down to 25 m s 5 Kn The greater the sensitivity the more the instrument will respond to speed variation However these instruments are usually fragile and pose reliability problems They are usually designed for a low speed range and can be damaged by gusty turbulent air 8 1 3 Wind Direction Tri axial hot film or ultra sonic sensors will provide the total vector in space wind speed and direction These high response instruments will show rapid changes For most applications a wind vane will provide acceptable data With careful at tention to design and construction details vanes can provide data accurate to 0 5 degrees 8 1 3 Ambient Air Temperature The temperature sensors are usually placed at the same Location on the tower as the wind instruments This eases the oquipment installation and aids the data correlation The sensor must be shielded from solar radiation
99. ry to only measure power input to the rotors With new engines or new installations it may be required to measure every element in the propulsion system The most direct method to determine power is measurement of torque and speed which are then used to calculate power Other methods include use of fuel flow and temperature in con junction with engine charts and engine characteristics data The engine airframe interface must be established in terms of inlet and exhaust characteristics Engine cooling and vibration can also have a significant impact on suitability When system losses must be determined each component will be instrumented to provide input and output data Accessory power must be determined for any power extracted to operate aircraft systems the instrumentation may include electrical hydraulic or pneumatic measurements For tests of the dynamic compatibility of new or modified engine airframe combinations and teste to evaluate engine rotor response characteristics the accuracy may need to be compromised to obtain satisfactory dynamic response from the instrumentation In some cases redundant instrumentation will be necessary to meet both steady state accuracy and dynamic response requirements 3 1 Shaft Speed Measurements Contained within the propulsion system are a wide variety of rotating components and measurements of the rotational velocities of these components are often of critical interest to the test being conducted Eva
100. s Technician The data systems technician has overall responsibility for completion maintain ing currency and distributing copies of the Recorded Instrument Parameter List He is specifically responsible for providing information in the CALIBRATION section after he has run the calibration and it has been approved by the engineers The specific procedures used to ccemplete the initial list follow ES OT ai 2 SR BRET ES ee te SE a 1 As calibration data sheets are received calibrations are run and plotted They are checked for obvious errors 2 The master plot and data sheet are then given to the I E and a plot copy to the F T E for approval 3 After approval or recalibration CALIBRATION information and TRANSDUCER information are inserted in the master list draft The calibration deck is assembled simultaneously 4 Master plots and Calibration data sheets are then filed in the aircraft instrumentation file in the I E office 5 After receipt and approval of the last calibration tne draft form is typed for the initial master list 6 A listing is made of the completed calibration deck The original is filed in the aircraft tostrunenta tinn file and a copy given to the F T E Document provided by SpaceAge Control nic thitp spaceagecontrol com oe ree T em E WEES SCS PRET REE Seay ME MERE DEET Daas aea ERS UNITS Engineering units used in calibrat
101. s uncaged and then will show deviations from initial Be aircraft heading The range is then 180 from that heading with a linearity of 0 5 of k i full range For some instrumente yawing more than 180 will cause the gyro to tumble ay n Si Certain tests of navigational systems or earth referenced maneuvers require magnetic headings This data is obtained by use of the aircraft gyroscope with signal conditioning such as a synchro to D C or binary converter which tracks the compass head ing Yaw attitude changes can be computed using Euler angle transformations Bare 4 2 Angular Rate Angular rate data can be calculated from the attitude gyroscope output In addi tion to the computation requirements a portion of the data is lost with this approach and helicopter flight tests usually require installation of rate gyroscopes The rate will be 4 measured from an initial test condition which can be either static or dynamic Following a i control input the maximum rate will usually occur within two seconds and appropriate sensor characteristics should be selected The rates are measured for each aircraft axis and the location of the sensor wust be accurately established The rates generated during the tests are considerably less than for high performance fixed wing aircraft and sensor k 1 H Eo 9 he ge RF selection should be influenced accordingly The range must be sufficient to encompass 4 i x aircraft m
102. station The sensors must have a sensitivity range and response compatible i with the data required For a large number of sensors it is common practice to use a j scani valve arrangement where each sensor is switched to single pressure transducer in sequence rather than being continually input to the transducer The time increment between samplings must be carefully considered A good technique is to measure selected sensors continuously while still including them in the sampling sequence This provides a check on i data validity and aids correlation of data from all the sensors Total pressure ranges from ambient to dynamic pressure at maximum airspeed Pressure should be measured very i accurately since a small change can result in significant differences in calculated engine power available or power required When scani valve arrangements are used dynamic response of the pressure transducer must be considered to insure proper performance with the multi piexed inputs Some inlets have filters particle separators and flow control or by pass devices which may require evaluation In most cases a single upstream and downstream pressure differential across the device will be adequate however a rake similar to that for the compressor face may be necessary to provide the needed data It is only possible to gener alize here and let specific decisions be made for exch individual situation i i d d Document provided by SpaceAge Contr
103. t A common approach is to have combinations of the variations mentioned above The user tests cannot be done with the quantitative accuracy that is possible in the performance or stability and control tests The greatest difference is in the atmospheric conditions Performance and stability tests are normally conducted in a stable air mass while the user tests are conducted in opera tional conditions Turbulence wind snow ice rain and dust are ever changing and create complex effects that are presently beyond our ability to account for or measure Thus the suitability of the machine is largely determined by the pilot comments or the capability to accomplish a specific task at a general set of conditions While inexact from an engineering viewpoint these tests are a good measure of the ability of the men to live with the machines and of the capability of the machine to accomplish the mission 1 3 Instrumentation Environment Helicopter instrumentation often must survive in conditions more adverse than are generally present during flight tests of fixed wing aircraft Smali helicopters have limited space available and various compartments may be used The instrumentation system may have components separated which can cause many lectrical problems Electrical power may be limited and in the case of transmission driven alternators power may be inter rupted at low rotor speeds Trroughout the helicopter high vibrations should be expected The
104. t conditions are a critical item in the analysis of the propulsion system and for determination of all aircraft performance Data must be obtained which will show the nature of the flow into the compressor establish the mass into the engine and establish the starting point for a thermodynamic analysis of the engine The inlet may be all or any part of the total ducting shaping guiding or holding apparatus between the free air stream and the compressor face Consideration must be given to the extreme range of conditions generated by the helicopter flight regime Vertical forward rearward and lateral flight produce the full range in terms of sideslip and angle of attack Rotor downwash is usually present and there may be engine exhaust gas ingestion caused by circu lation of the rotor wash The inlet performance is usually defined in terms of pressure and temperature conditions at the engine compressor face Test requirements may dictate establishing the turbulence or distortion in the inlet flow During the instrumentation design phase special care should be given to obtaining data compatible with any previous umant provided 9y ed ures Sono no http spaceagecontrol com AEAN ts Fei an Tis a da eee SE ae Nee T EESE engine calibrations or any data needed to operate the engine computer program Any inatru mentation must be fully certified for the expected dynamic pressures temperatures and vibrations before it is placed in the
105. t of occupants A similar procedure is used for avionics and cargo compartments The quantity of aivilow and heat being provided to the compartment is measured at fr the duct outlet or the heater Outlet air temperature is measured with a thermocouple as 1 r discussed aboye Total and static pressure sensors are also placed at the outlet to deter i p Y y Ag we i mine airflow Selection of sensors must consider the very low velocities to be expected Mf Planning information can be obtained from design specifications or from systems test TA results Humidity in the compartment can be measured with any suitable hygrometer The airflow patterns within the compartment can be measured with a hot wire anemometer ihe anemometers can be mounted on a rack and moved to different locations or the sensors can be placed at the position where the temperature profile is being determined Air quality can be measured with various instruments to monitor and sample dif ferent types of gases and toxicity levels In addition it is common practice to obtain air samples in suitable containers and then to perform a laboratory analysis 4 7 2 Surface Temperature 4 The temperatures of interior compartment surfaces and any exposed ducting are measured with thermocouples Calculation of heat logs through windows requires that both interior and exterior surface temperatures be measured sie eee ot SISTENT NTS to Tes 7
106. tank level devices must also consider aircraft attitudes and flight conditions 3 7 1 Fuel Flow The fuel mass being used by the engine can be measured directly with a mass flowmeter or calculated from volume and temperature measurements Sensors are available for various flow rates and the appropriate one should be selecte for each engine instal lation It should be noted that small helicopters have flow rate 3 low as 9 kg hr 20 lb hr at idle while large machines use up to 1130 kg hr 2500 lb hr at rated power In flight test the flow is most commonly determined by measuring volume flow and temperature and then calculating the mass flow The volume flow is obtained from a turbine flowmeter Prior to installation the flowmeter is calibrated so that for a given turbine rotational speed the flow can be determined Calibration of the flowmeter is critical and requires precise control of calibration fuel parameters to insure definition of the number of cycles of output from the turbine flowmeter for a specific unit of flow Most test turbine fuel systems use the cyclic output to determine total fuel used information Both cockpit display and digital output for instrumentation system recording of fuel total are generally used Fuel flow rate is computed by interfacing a frequency to D C voltage level converter to the turbine output Again both cockpit and instrumentation system recording of this data is normally used The flowmeter creates a small
107. tcbility an control or user data cun be obtained during the performance tests Stability and contro ests Document provided by SpaceAge Control Inc hitpillspacganecon ral casiMlatioaaasall 8 SORES PSE ee ca Heft US HTP RP eT OS RES E re er PERSE AER Eee oR MRE INN eee ot en are a combined quantitative and qualitative effort Fur these tests emphasis is placed on flight control systems aircraft motions and positions Power and atmospheric conditions are not as critical as for th performaice tests The data provide design information and establish flight capability and flight envelopes Qualitative pilot comments are used to assess pilot workload and man machine compatibility Test pilots must relate their experi ences with the test vehicle to the expected ability of the operational pilots A very important part of these tests is the failure mode tests Characteristics of the control system are evaluated in great detail ard all possible combinations of failures are con sidered Appropriate caution or warning notes are generated and placed in the pilots operating manuals User tests will be peculiar to the mission of the organization or dictated by the aircraft characteristics Those tests may be quantitative or qualitative Operators can be either test pilots or user pilots The instrumentation may be special test equipment or it can be the standard aircraft equipmen
108. te electrical circuits must be developed to provide signals to cockpit indicators and aircraft systems These circuits are normally used as input to the instrumentation system and care must be taken not to alter the operation of the standard torquemeter system Isolation amplifiers may be required to insure separation of the aircraft torque system and the instrumentation system In most cases the signal level of the torque system will be less than one volt and noise reduction techniques should be included The engine is placed in a test cell and the torque is measured directly with a dynamometer and the indicator reading is noted From this calibration torque can be determined for any indicated reading Engine torquemeters have an accuracy on the order of t5 although in one instance an accuracy of t1 is cluimed he engine torquemeter output must be recorded during the test since the operators manual will be developed in terms of the power indication to the pilot Vuring development of new engines or for standard torquemeters that provide inadequate data it may be necessary to install a test torque measuring system The test power measurement system is usually placed on the engine output shaft The torque must be measured on the shaft for which the speed measurement was taken Extreme care and close coordination with the flight test engineer is needed to determine what power is being measured and that it is the correct power for the data requiremen
109. temperature The transmitters are simultaneously pulsed and at a later time typically 200 to 300 microseconds the wave arrives at the receivers EEP PEST METEU AOE E PER EDA ee Se e T By writing the equations for the other two transmitter receiver pairs as a function of their respective transit times three equations with the three unknown vectors result There are not of a closed form because the vectors are a function of total vector and therefore must be solved by iterative or feedback methods With zero relative wind velocity the three transit times will be identical and equal to the ultrasonic wave transit time at the particular temperature With a relative wind along the X axis only the times are equal but increase in value for forward aircraft motion For relative wind in an arbitrary direction the three times will be different in value In general the times can be considered as quantities which vary by a percentage around the still air value The total vector can be used to provide three airspeed components as well as angles of attack and sideslip J PROPULSION SYSTEM The propulsion system data is critical for ail performance tests The system includes engines transmissions and drive train components Emphasis is placed on para meters which are used to determine power required and power available Power measurements vary considerably with test objectives For engines which have been previously defined it may be necessa
110. thout electrical mechanical or explosive deficien cies The flight test of a weapon system then becomes a task of determining the weapon compatibility with the helicopter and crew The weapon system will usually add drag introduce loads into the airframe and alter the stability and control characteristics i i B 7 1 Forces and Motions Externally mounted weapons generate drag whigh transmits forces through the attachment hardware into the airframe Recoil forces will be added during firing Strain i gages are placed on the airframe or on the weapon structure attached to the airframe In l the non firing mode tho weapon will react to rotor induced vibration through the airframe and when firing vibrations will be generated by the weapon Instrumentation design must i consider that rotor vibrations are usually low frequency while weapon firing rates can be F up to 2000 rounds per minute which generates high frequency reactions Accelerometers are usually oriented along the recoil axis of the weapon Traversing or elevating weapons may require accelerometers in three axes ce engage Mee E GT SS E EE P 7 2 Firing Effects Gun type weapons generate significant overpressures which can cause structural damage A pressure transducer is attached to the airframe where air pressures are the i greatest The number and location of sensors will depend on each particular installation Missiles create exhausts which can impi
111. tion Aeroflex True Airspeed Vector System Low Airspeed Systam Final Report USAASTA Report No 71 30 2 P 2 4 17 F Dominick K R Ferrell Cpt J O Conner US Army Aviation Systems Test Activity 1975 Flight Evaluation Elliott Dual Axis Low Airspeed System LASSIE II Low Airspeed Sensor Final Report VI USAASTA Final Report 71 30 6 P 12 18 18 AGARD No 219 Range Instrumentation Weapons Systems Testing and Related Techniques 19 W Abbott Cpt S Spring Maj R Stewart US Army Aviation Engineering Flight Activity 1977 Flight Evaluation J TEC VT 1003 Vector Airspeed Sensing System Final Report USAAEFA Report No 75 17 2 P 10 13 TESTI oe meee a 20 W Abbott B Boirun Cpt G Hill Cpt J Tavares US Army Aviation Engineering Flight Activity 1977 Flight Evaluation Pacer Systems Low Range Airspeed System LORAS 1000 Final Report USAAEFA Report No 75 17 1 P 11 21 Ber ree 21 W Abbott Maj J Guin US Army Aviation Engineering Flight Activity 1977 Flight Evaluation Rosemount Low Range Orthogonal Airspeed System with 853G Sensor Final Report 75 17 3 P 12 17 http spaceagecontrol com easi Document provided by SpaceAge Control Inc tate eset SABE AA PELE RR ts els ea aad sk ia As ia WE ET any SERRE tA a i ald 22 B Boirun Cpt G Hill CW3 J Miess US Army Aviation Engineering Flight Activity 1976 Flight Evaluation Honeywell Ultrasonic Wind Vector Sensor Syste
112. ts Note should be made of the power being measured relative to transmissions and power extraction sources Resistance type strain gages are commonly used to sense the torsion in the shaft Tempera ture compensation must be adequate for the installation and consideration must be given to shaft bending moments he strain gages are connected to a slip ring brush assembly which transmits the signal The electrical and mechanical properties of the slip ring assembly must be compat ible with the strain gages being used After the strain gage installation the shaft is calibrated in terms of force and deflection Checks are made for adequacy of compensation efforts and if necessary those influences are included in the calibration It may also be necessary to dynamically balance the shaft to compensate for the added instrumentation In some cases slip rings have been replaced with telemetering systems which transmit torque data from the rotating shaft to a stationary receiver 3 3 Shaft Torque Torque measurements on the individual drive train shafts are made at the same place as the speed measurements Shaft speeds diameters and environments will vary considerably and special care must be taken to compensate for mechanical or environmental effects A strain gage system similar to the engine torque is most commonly used Me chanical optical and electromagnetic systems have been used to measure shaft deflection with applied torque 3 4 Inlet Inle
113. tt The Elliott low airspeed system is manufactured by Elliott Flight Automation Ltd Airport Works Rochester Kent England In the United States the equipment is the respon sibility of E A Industrial Corporation Chamblee Georgia the associate company The sensor and vector resolution is shown in Figure 2 2 2 2 Document provided by SpaceAge Control Inc http spaceagecontrol com PPP acs eae o hied AA l kaataki air at itae a i ja eae ee i an eee we SPS po airs ie cea En Re ES pe Se wom PRESSURE 7 STATIC PRESSURE CIRCULAR VANE SENSEO LATERAL AIRSPEED V sin 8 SENSED FORWARD AIRSPEED aincrart FORE AFT i 76 i ores AIRCRAFT LATERAL PROBE Figure 2 2 2 2 Elliott Airspeed Sensor The system includes a swiveling total and static pressure sensing probe a computer and airspeed indicators for three axes The resultant downwash aligns the probe with local relative wind vector sum of aircraft velocity and rotor induced velocity and provides adequate dynamic pressure at all airspeeds The angle of the probe and the differential pressure are used to calculate aircraft speed and relative wind direction Static pressure is measured and rate of change is calculated to provide rate of climb information The airspeeds presented to the pilot are longitudinal l
114. ttack and sideslip With the doppler system four signals are transmitted and speed is determined from the doppler shift in the return signal Inertial systema use accelero meters in the aircraft axes to provide data for speed calculations Depending on the system it may be desirable to record either the accelerometer data or the differentiated output Navigation systems are intended for trimmed flight and the true airspeed calcu lations may be affected by sideslip angle This ig of particular importance in helicopters which frequently have large sideslip angles at low speed or during crosswind maneuvers Complete space positioning data ia ootained from the navigation system ground speed com ponents combined with a low airspeed omni directional airspeed system and ground atmos pheric measurements Document provided by SpaceAge Control Inc http spaceagecontrol com sezanar arnee ebshedes vert bidet setters deri aiedin asa L s pe bjar tent acts the PES eS al REFERENCES 1 A Pool and D Bosman AGARD AG 160 Vol 1 Basic Principles of Flight Test Instrumen tation Engineering 2 D W Blincow Prepared under AFC Contract AT 04 3 Nuclear Helicopter Lift Indicator NUH ELI 1970 SAN 4007 1 P 56 3 C R Duke B Y Cho D E Cressman Prepared for Naval Development Center Warminster PA under contract N61169 69 C 0578 1970 Development of a Feasibility Model Air Density Gauge Final R
115. um element prebe designed for immersion in fluids Any number of platinum element probes can be obtained for this application but attention must be given tc proper design of the installation utiiizi g them Frictional heating of the probe due to fluid motion stem conduction errors and flow obstruction must be considered in the design Some relief from these requirements for engineering can be obtained by implementing a platinum probe tubing combination sometimes called an in line sensor These sensors install as a short piece of tubing with the platinum probe integrated into the tubing by the sensor vendor and can be ordered to meet the specific application at uand 3 7 3 Fuel Quantity Prior to flight the fuel mass and aircraft weight are precisely known There fore an accurate in flight gross weight requires determination of fuel used from engine start to the time test data is recorded Fuel us d is determined by the flowmeter cycles and the fuel density in the fuel tank Specific gravity of the fuel in the tank is estab lished in the laboratory for samples iaken before and after each flight Any flow through fuel return or by pass lines is measured and used to correct fuel flow and fuel used data 3 8 Power Extraction Power may be extracted from the engine to operate electrical systems hydraulic systems and environmental control systems for the occupants For exacting aircra t performance tests these powers must be considered i
116. us the component of the airstream parallel to the turbine axis is synchronous with the turbine speed OPTICAL PULSE GENERATOR TURBINE ORIVE MOTOR FLOW STRAIGHTENERS SPEED SENSING BOLOMETER DIRECTION SENSING BOLOMETER JTE TURBINE fi ABAD BOVABer Basse SALT eee TO ee Bee eS NE Cd iy 9 E f a ME E A A bm a a a m a a m a Poa 1d A A helo eX e a DAA OIREAN ERAMEEERAMNNER 4 en EE LE Re y envn SS Stree rrr tt 1 RRB BSSRwm ith DIRECTION SYNCHRO AC MOTOR DIRECTION ORIVE Figure 2 2 2 1 Aeroflex Airspeed Sensor The airspeed and direction sensors drive a visual indicator and provide DC volt age outputs The indicator contains a roller suspended servo driven tape marked in m s 1 Kn increments at its center to display airspeed in the range of zero to 180 m s 350 Kn At the perimeter of the indicator face a servo driven ring continuously displays the sensor head position relative to the sensor base through 360 degrees of rotation The DC output has three separate output recording terminals Each output consists of four buffered channels and can drive as many as four oscillograph galvanometers or similar recorders The DC outputs consist of a coarse signal for airspeed zero to 130 m s 250 Kn a coarse signal for direction zero to 360 degrees and a fine signal for airspeed and direction which cycles every 13 m s 25 Kn and 36 degrees respectively b Ellio
117. ut is 1 G The angle ig calculated from the measured G recorded The accelerometer output is also displayed on the pilot s instrument panel to assist in estab lishing a hover that minimizes the deviation angle 8 4 Space Positioning Many tests require precise measurement of the helicopter position in space at a given time These tests include take off landing acceleration in three axes and various agility maneuvers In all cases the distances are relatively short and the aircraft is near the ground The space positioning system includes data acquisition range support and atmospheric measurements and provisions for subsequent data processing The initial instrumentation planning should consider 1 Test Site whether the tests will be conducted on an instrumented rangs or at a remote site 2 Equipment location whether the equipment will be located in the aircraft or on the ground 3 data recording whether the data will be ground recorded or recorded on the tect aircraft In all cases provision must be made for items nevded to control and conduct the test as well as document the data for later correlation merging and processing With rare exceptions the space positioning system ground station layout will be needed This will normally be precisely determined and readily available from an instru mented range For temporary installations a survey is required The space position stens provide motion relative to the ground whil
118. vision should be made for automatic data numbering and data avent markers Event markers ave extremely important for the flight crew to note significant data points or unusual occurrences during the test Cockpit and or ground playback and monitoring capability contribuvec to data validity and assurance that desired test data is being recorded When feasible and cost effsctive the data should be machine processed In most modern faci tities the instrumentation and data processing systems are difficult to sepa rate Thus it is mandatory to consider this interface when designing the instrumentation system 1 5 Instailation The instrumentation installation must be designed to be compatible with the test vehicle acilitate pre flight inspection and maintenance and to minimize crow workload during tha testing Access to the test vebicle ur scale drawings are nec2ssary to establish the location of instrumentation The instrumentation buildup can usually be accomplished more easily in the shop than in the aircraft The instrumentaticn layout must consider a Accessibility for check out and mainterance b Structural integrity for flight safety and crash worthiness c Vass locations for aircraft weight and balance considerations d Possible inflnence on vibration characteristics e Convenience for flight crew operation An evfective test program requires that the pre and post flight instrumentation activities can be accomplished in a minimu
119. will be transmitted by attaching a draft form with the first 9 columns except PCM FM CHAN completed This will be the basis for completing the master list Polarities should be conventional with the possible exception of vertical acceleration which is positive downward in some systems Use of these polarities is mandatory Provisions should be made for multi engine rotor helicopters and FM data Nominal ranges and gains zero count and max count or per count are used on the draft Actual ranges and gains will be obtained from the calibrations Instrumentation Engineer I E The instrumentation engineer will arrange for calibration and installation of the requested parameters In coordination with the project engineer or programmer he will complete PCM FM CHANNEL assignments and SIGNAL CONDITIONER information The instrumentation engineer and flight test engineer will review and approve all calibrations after they have been nere and plotted Calibrations are then provided to the instrumentation technician Instrumentation Technician 1 T The instrumentation technician is responsible for the physical installation of all requested parameters in coordination with the flight test and instrumentation engi neers He is also responsible for performing required on board calibrations The I T will provide all information required in the transducer section for each parameter This will generally be done via the calibration data sheets Data System
120. y which unbalances the error sensing bridge A servo system then rotates the pylon until the Document provided by SpaceAge G eh te lh eee soe saad ea rath ae iol 5 MONE et ae ee EEE EE SLI LO BEET ET ERIE DED STS i 3 f E RAUR ni ESTORSER TIRARE Ne es ere ST NAR 9 bridge is balanced and the stream tube is parallel with the airflow A vertical stabilizer is externally mounted on the aft portion of the stream tube to provide directional sta bility at high speeds The true airspeed sensing unit is a hollow tube mounted on a pylon base The forward portion of the tube contains a honeycomb structure which assures axial flow at the inlet and also creates turbulent flow through the stream tube throughout the speed range of the sengor The rear portion of the tube contains a 18 blade turbine and two V bolometer assemblies aft of the turbine The principle of operation is based on the premise that for a given airflow through the turbine duct the turbine can be rotated at a speed synchronous speed which will permit undisturbed axial flow An illustration of the sensor and a vector representation of its operation is shown in Figure 2 2 2 1 If the turbine is not at synchronous speed with the airstream V R Vrp the two V bolometer assemblies sense the resultant airflow V as a deviatio from xial flow and a servo motor adjusts the turbine speed until the error signal is nulled Th
121. ystems within 5 m s 10 Kn at 20 m s 4000 ft min vertical rate 2 2 2 Omni Directional Airspeed Systems Many helicopter tests require airspeed information at low airspeeds and in various directions Several systems have been developed which provide data in hover vertical climb and descent and during sideward or rearward flight These systems are also operable in high speed conventional flight Hover performance is very sensitive to relative wind which must be measured within 0 5 m s 1 Kn The wind direction can also affect the power required or critical directional control margin and should be measured with an accuracy of 2 degrees Location of the sensor is critical since it is desired to measure aircraft velocity and not local flow conditions Rotor wash is the largest single factor although disturbed flow from the fuselage wings or stores must also be considered It is expected that each installation on 2 particular aircraft model will be unique and the system will require a flight calibration to determine the position error Typical changes in position error with sensor location are shown by Ref 15 Most of the low airspeed systems have been developed further since the referenced tests were completed In fairness to all manufac turers and to avoid misinforming the reader resolution threshold and accuracy numbers will not be presented here Capabilities of the various systems are summarized in Table 2 2 2 1 Performance and special c

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