Model-Aircraft Instrumentation SystemDavidMedina

EscueladeIngenierıasE3TIngenierıaElectronica

GrupoERAUniversidadIndustrialdeSantander

Bucaramanga,ColombiaEmail:david.medina1@correo.uis.edu.co

NicolasBeltranEscueladeIngenierıasE3T

IngenierıaElectronicaGrupoERA

UniversidadIndustrialdeSantanderBucaramanga,Colombia

Email:nicolas.beltran@correo.uis.edu.co

YesidBarreraEscueladeIngenierıasE3T

IngenierıaElectronicaGrupoERA

UniversidadIndustrialdeSantanderBucaramanga,Colombia

Email:yesid.barrera@correo.uis.edu.co

Abstract—To know how to measure physics magnitudes usingelectronic devices is very important nowadays and it is betterif that knowledge is acquired in the undergraduate level. Oneinteresting measuring process is developed in aviation, where itis important to know the value of many variables to get a softflight and a good behavior on air.One easy way to get in touch with the aviation field is throughmodel aircrafts. For this reason this paper shows the process ofdesign and implementation of a system to measure speed, height,motor temperature and RPM1 of the model aircraft Ugly-Stick40, and to transmmit them to be displayed on ground.

Index Terms—OpAmp, Differential pressure, Thermocouple,Microcontroller

I. INTRODUCTION

This project was develop by students of 8th semester ofElectronic Engineering, looking foward to apply the knowl-edge acquired in the Analog Systems Design subject to processand measure physics magnitudes common in the aeronauticfield. Through this paper, it is shown how the system wasdeveloped, the decision process in the design and the im-plementation of the measuring system into the model aircraftUgly-Stick 40. The system has a modular design to improvethe design time through tasks distribution. In each modulethere are three stages: signal acquiring, signal conditioning,and signal processing module. Additionally, each acquiremodule description has noise prediction. On the other hand itwas designed and implemented a RF link to transmit velocity,temperature, RPM and height data from the airplane to acomputer with a LabView GUI2 on ground.

II. GENERAL PROBLEM

First, it is necessary to measure motor temperature, modelaircraft speed against air, height against the pilot position andmotor RPM. Second, the measured data must be transmittedfrom the model aircraft to the pilot on ground, and it mustbe displayed in a Labview GUI, so that he can take decisionsto preserve the integrity of the airplane. All of this must beconsidered for Colombian atmosphere.

978-1-4244-6742-6/10/$26.00 c© 2010 IEEE

1Revolutions Per Minute2Graphic User Interface

III. DESIGN REQUIREMENTS

Collecting information from datasheets, experience andtests, the following paramaters are established:

• Maximum motor temperature: “200C”.• Minimum motor temperature: “-10C”.• Maximum aircraft speed: “150 km/h”.• Maximum flying height: “3500 m.a.s.l.3”.• Maximum RPM: “24000 rpm”.• RF link range: “1 km”.• Graphical data view.

IV. SOLUTION PROPOSAL

A. Temperature SensingAn overheating limit of “300C” is supposed. IC4 sensors

are not a good option, taking into account the motor tempera-ture’s variation range. For this reason, a thermocouple K withmeasure range from “−200C” to “600C”; It is great for ourworking range; more over standards thermocouple are cheapand easy to buy anywhere.

A thermocouple provides output voltage in the order ofMicro-Volts, so it’s necessary to amplify it to take it intodetectable values by the ADC5 module of a µC 6. This voltagesignal must also be adequated.

Another problem that a thermocouple involves is the de-compensation at the cold junction. To solve this problem wepropose compensation through software, because it is cheaperthan buying an IC compensator.

1) Software Compensation: The cold junction will be ex-posed to ambient temperature variation. This makes veryimportant to assure a dynamic compensation.

Using a temperature analogical sensor IC, the junctiontemperature is measured in order to compensate by softwareinside the µC through the eq.1

Tm = Ttc + Tcj (1)

Where Tm is the motor temperature to measure, Ttc is thetemperature measured with the thermocouple and Tcj is thecold junction temperature.

3Meters above sea level4Integrated Circuit5Analog to Digital Converter6Micro- Controller

The IC selected to measure cold juntion temperature isthe STLM20 made by STmicroelectronics. Some of itsspecifications are shown in table IV-A1.

Table IPRINCIPAL CHARACTERISTICS OF STLM20

Characteristic ValueVS “2.4 - 5.5 V”Iq “8 µA”Zo “180 Ω”

Accuracy “±1.5 C”

The output voltage behavior of this IC between −10C and60C is described by eq.2.

Vo = −11.7mV/C ∗ T + 1.8641V (2)

Since the output signal is analogical, it must be passed troughthe generic ADC module of the µC. The generic ADC moduleinside µC can be configured as 8,10 or 12 bits resolution,which is more than output voltage sensitivity of STLM20. Forthis reason we only care about coupling impedance using abuffer opamp.

The characteristics of the OpAmp selected as Buffer areshown in table II

2) Thermocouple Signal adequacy: The use of one INA7 isunnecessary since common mode electrical noise has not hightlevel in this application because model aircrafts are flown inopen spaces. The topology for capturing and adequating thesignal provided by thermocouple is shown in figure 1 This

Figure 1. Thermocouple Adequacy System

topology uses OpAmps with high CMMR8, low internal noiseand low current consumption as showm in tables III and II.Looking at Fig. 1 the output voltage can be expressed as shownin Eq.3. If values exposed in the Fig.1 are remplaced, the Eq.4 can be obtained

Vo = (1 +R3

R2)(0.041 ∗ 10−3)Ttc “V ” (3)

Vo = 9.061 ∗ Ttc “mV ” (4)

Finally a passive Band-pass filter is used to select thedesired signal bandwidth. In this case we use a first orderfilter with cut off frequency of “22 Hz” approximately, tryingto make it has lower has possible with components without

7Instrumentation Amplifier8Common Mode Rejection Ratio

Table IIOPAMP TLV342A

Characteristic ValueVS “1.8 - 5 V”Iq “70 µA”

Offset “1.7 mV”GBw “2.3 MHz”

CMMR “95 dB”

Table IIIOPAMP TLV2721

Characteristic ValueVS “2.7 − 10 V”Iq “100 µA”VIO “650 “µV”IIO “60 pA”GBw “480 kHz”

CMMR “82 dB”

an very high value. This is made because motor temperaturedoes not change at high rates, so it can be assumed mainly asa DC component and it does not affect the system behavior.

B. Relative Speed Sensing

Model-airplane’s relative speed is very important to keepa stable flight and to assure stall speed. Using a Pitot-StaticTube is one of the most popular ways to measure speed.

The main opening in the pitot-static tube is placed in thesame direction of air’s flux. As the pitot tube has no exit, air’sflux get stucked and it generates a stucked pressure known astotal pressure. The lateral openings in the pitot-static tube areorthogonal to the air flux so they detect the static pressure.Then, the model aircraft’s speed can be calculated by usingEq.5 and Eq. 6.

PT = PS + PD (5)

V =

√2 ∗ PD

ρ(6)

Where PT is the total pressure, PS is the static pressure,PD is the dynamic pressure, V is the model-aircraft’s speedand ρ is the air density.

The stall speed is calculated taking all the model-airplane’sweight as a Wing Load and it’s pressure over the wing areaas the dynamic pressure in Eq.6. Then taking the standardweight and wing area of an Ugly Stick 409 its stall speed canbe calculated and its equal to “31.536 km/h”.

The Model-airplane’s relative speed is also calculated basedon Bernoulli equation. To do this, the difference between totalpressure and static pressure is measured with the differentialpressure sensor MP3V5004DP10, which can detect as far as“3.92 kPa”. This sensor has an output voltage described bythe Eq.7.

VOUT = VS ∗ ((0.2 ∗ PD) + 0.2) ± 2.5%VFss (7)

Where VOUT is the output voltage, VS is the supply voltageof the sensor, PD is the differential presssure in “kPa” andVFss is an standard reference value equal to “1.8 V”.

The output signal from the sensor is then filtered, as donewith the thermocouple, and amplified so the output voltagecorresponding to “120 km/h” fits with the supply voltage ofthe ADC module in the microcontroller. With these adecuatedsignal a better use of the dynamic range and a better high

9Instruction manual with standar values available athttp://manuals.hobbico.com/gpm/gpma1220-1221-manual.pdf

10made by Freescale Semiconductors

frecuency noise are obtained. Finally, the amplified signalgoes through a buffer op-amp to get impedance coupling. Thecircuit can be seen at Fig.2

Figure 2. Differential Pressure Adequacy System

C. Relative Height Sensing

This application was designed to measure a maximumrelative height of “1 km” in heights from “0 m.a.s.l” to “3000m.a.s.l”, because the radio control scope is “1 km” and theplaces where the model aircraft is going to fly can have thisheight.

The absolute flying height of the aircraft is measured byusing the relationship between height over sea level andabsolute atmospheric pressure, thought the definition of astandard atmosphere11. An initial height is calculated at thecircuit starting process, so the relative flying height could bereferenced against it inside the µC.

The absolute pressure sensing is made with an absolutepressure sensor, reference MP3H6115A6U with a measuringrange from “15 kPa” to “115 kPa”. The sensor transforms theinput pressure into output voltage as given in eq.8. To calculatethe expression of output voltage Vs = “3.3V ”, maximumambient temperature of 40C and information provided by themanufacturer12 are used.

Vo = (0.0297 ∗ Pabs − 0.3135) ± 0.04455 “V ” (8)

In order to calculate the height referenced to the sea level theeq.9 is used, with Pref = “101.325kPa”, which is the valueof atmospheric pressure at “0 m.a.s.l.” according the standardatmosphere and P is the value of pressure measured with thesensor. The determination of the signal adequacy depends onthe ADC resolution and output voltage range. To know theoutput voltage variations within the height range the table IV-Cis analyzed.

h =288.15(1 − (P/Pref )0.19026)

0.0065“m.a.s.l” (9)

Based on the information of table IV-C, the range of outputvoltage goes from “1.5 V” to “2.7 V” and the minimumvariation of output voltage is “0.25144 V” when it varies from

11referenciar12Freescale Semiconductor

Table IVPRESSURE VARIATION IN INTEREST RANGE OF M.A.S.L

Height Vo“0 m.a.s.l” “2.69585 V”“1 m.a.s.l” “2.69650 V”

“999 m.a.s.l” “2.35609 V”“1000 m.a.s.l” “2.35577 V”“3499 m.a.s.l” “1.63993 V”“3500 m.a.s.l” “1.63968 V”

“3499 m.a.s.l” to “3500 m.a.s.l”.Now, to know how the ADC resolution affects the measuringprocess, the sensitivity of the ADC is evaluated with resolu-tions given for 8,10 and 12 bits and a reference supply voltage“3.3 V”. The information obtained is shown in table IV-C.

Table VADC RESOLUTION DEPENDING OF NUMBER OF BITS

Bits Resolution8 “12.94118 mV” “51.46881 m”

10 “3.22581 mV” “12.82947 m”12 “0.80586 mV” “3.22343 m”

By analising the table IV-C it can be seen that getting thesignal directly from the sensor to the ADC module is hasmaximum mean error of “±3.22 m.a.s.l.” with the 12 bitsADC. It’s possible to improve the measuring resolution withextra stages. However the error obtained without these stages isacceptable and a µC with 12 bit ADC module will be selected.

As known, measured values based on standard atmospherewould be wrong, for this reason the value measured is compen-sated inside the µC by using the charles’s law for ideal gasesapplied to two columns of air with same base and differentheight it’s possible to write the eq.10

hc =TaTsho (10)

Where hc is the Compensated Height, Ta is the ambienttemperature, Ts is the expected temperature at the measuredheight and ho is measured height.

Finally the fixed bandwidth is gotten through a filter withcut-off frequency of “22 Hz” and the impedance coupling ismade through a buffer. The schematic for this stage is shownin figure 3.

Figure 3. Schematic view of stage sensing height

D. Engine RPM sensing

Engine RPM are associated to the displacement of air fromthe propeller, and thus to the push that engine generates overthe entire model aircraft.

The Diferential Peak-Detecting Sensor IC, with referenceATS682LSH made by Allegro Microsystems Inc, was selectedas the sensing element to measure RPM of the engine. It’sassembled to the system to detect the passage of the counter-weight of the crankshaft, and thus each spin of the engine.

This integrated circuit incorporates a dual-element Hall-effect circuit and signal processing that switches in responseto differential magnetic signals created by ferrous elementsthat affect the magnetic field generated by its inner magnet.

The IC was selected because of its reliability and ability toreduce magnet and system offsets and to calibrate the gain forair gap independent switchpoints.

Some of the IC caractheristics that made it ideal to ouraplications are:

• Supply Voltage: “4 V” - “24 V”• Operating Frecuency: “0 Hz” - “8 kHz”• Maximum Current: “16.8 mA”• Operative Temperature: “-40 C” to “150 C”

As the Output voltage in the GND pin of the IC has anoffset voltaje and it isn’t fixed to 0-3V range it is necessaryto use a comparator as shown in figure 4.

Figure 4. Schematic view of stage sensing rpm

The pulsing signal obtained from the circuit has the samefrequency that the engine crankshaft and, as consequence,that the propeller. This frequency is measured by detectingtwo following changes from a logic one to a logic zero andcounting the period between them with the Timer module ofthe microcontroller.

E. RF Link

The RF link is made of two main parts: the transmittermodule, and the receptor module. The first one recieves datafrom the µC and transmittes it. The second module receivesthe trasmitted data, demodulates it and then gives it to a toUSB13 IC so data could be get into a computer in a directway.

13Universal Serial Bus

1) Hardware and Considerations: In this system the bau-drate can be considered under “10 kbits/s”, as the measuredvariables have a slower rate of change. It will be used abaudrate of “1200 bits/s”, what gives a better communication,a longer reception area, and more noise inmmunity by keepingthe spectra of transmitted data closer to the carrier. These“1200 bits/s” are enought because it’s needed to transmite fourvariables using about “20 bits” for each one.

The carrier selected for the RF link is at about “418 MHz”,which is far from the radio-controller carrier that is at “72,53Mhz”. As it is necessary to assure a “1km” transmittingdistance the Eq.11 and the Eq.12 are used.

PTx − LCTx+GTx − La +GRx − LCRx

= M + S (11)La = 32.4 + 20 ∗ logF + 20 ∗ logD (12)

Where, PTx is the Transmitter output Power “dBm”, LCTxare

the antenna connector losses at Tx “dB”, GTx is the antennagain at Tx “dBi”, La are the electromagnetic signal losses inthe air “dB”, GRx is the antenna gain at Rx “dBi”, LCRx

arethe antenna connector losses al Rx “dB”, M is the securitymargin for the RF link “dB”, S is the receptor sensitivity“dBm”, F is the carrier frequency “MHz” and D is the RFline of sight distance “km”.

The RF modules TXM-418-LR and RXM-428-LR of theLR series of Linx Technologies are selected for the RFlink because of their low cost, reliability, support and goodperformance in open spaces. They also require the use onan antenna connector SMA and a antenna ANT-418-CW-HW.This RF link has the following characteristics:PTx = “4dBm”, LCTx

= LCRx= “0.25dB”, GTx = GRx =

“0dBi”, M = “30dB”, S = “ − 112dBm”With these values, and using Eq.11 it can be calculated that

La = “85.5dB”. With values of La and F = “418MHz”were replaced in Eq.12 was obtained D = “1.08 km” asrequired by the system.

Due to the module’s ease of integration, just a few con-nections for a effective data transferring are required. Nooscillator, capacitor or arrays of resistances are required. Thedata transferring process from the µC to the transmitter isdone through the SCI module adjusted at a baud-rate of “1200bits/s”. A picture of the transmitter an receptor module isshown at Fig.5.

Figure 5. Images of RF implementation

2) Data Transmition: The information transmitted shouldhave a standard protocol to determine the value of each vari-able and if the received data is the same as those transmitted,as they could become noisy and unpredictable due to wirelesscommunication.

In order to make it as ease as possible, the transmittedinformation has the format shown in Eq.13. In this format, aninitial letter identifies which variable is being sent, followedby five numbers in ASCII code, giving the variable value, anda comma as a finishing indicator. When all four variables havebeen sent a “\n\r” is sent to start a next line.

T#####, R#####, V#####, A#####\n\r(13)

This transmition is made as redundant as posible to comparereceived data and detect mismatches.

F. Noise Analysis

To calculus of the noise contribution at the system isdivided in stages. It’is considered than the pressure sensorsand the temperature sensor work with frequencies until “10Hz”, allowing just a contribution of Johnson-Nyquist noise.The spectral power density of this noise is given by eq. 14and the output noise is given by Eq.15

V 2n = 4KBTRG (14)

Vn = G√V 2n ∆f (15)

For the absolute pressure sensor the design only has oneresistance of “22 kΩ” and one capacitor of “0.33 µF” thatcontribute to the noise in a bandwidth of “∆f=10 Hz”.

For the stages of signal adequacy RG is the equivalentresistance at OpAmp’s input and G is the OpAmp gain.

It has been calculated the noise provided by external ele-ments, but the internal noise contributions are neglected be-cause the OpAmps have an internal voltage and current noisewith values less than 1% of the noise provided by externalelements, according to the manufacturer, for the bandwidthused in this application.

Table IV-F shows the total noise introduced by the system.

Table VIESTIMATE NOISE FOR THE SYSTEM

Sensor Total NoiseHeight “60.3589 nV”Speed “0.2183 µV”

Temperature “2.8381 µV”Total noise

of the System “3.1167 µV”

G. Graphical View

An easy to use graphical user interface is necessary tofacilitate the interaction between the pilot and the flyingvariables of the model aircraft.

LabView 8.6 was used to develop an dynamic GUI to showthe actual value of the flying variables. It is able to receive the

information trough an emulated COM14 serial port, compare itwith the last packages received and extract the variable actualvalues from it.

The functional blocks used to configure the COM port anreceive information from it are VISA Configure Serial port VIand VISA Read Function, correspondingly.

Using the Match Pattern function of the String librarythe received data can be segmented in the different variablevalues. This information segments are in ASCII format so aDecimal String To Number block is used to get the integercorresponding to each variable.

Finally, using some graphic indicators the different valuesare displayed in a friendly GUI as shown in Fig.6

Figure 6. Main Window of the GUI developed with LabView 8.6

V. MICROCONTROLLER SELECTION

To achieve the objectives was used a µC easily pro-grammable, with C language support, ability of float pointcalculation, serial interface, 12 bit ADC, SM15 package andavailability of pins.The µC MC9S08QE32 of Freescale Semiconductors was se-lected because the in-circuit programmer was available, it washad experience with the core and meets specification amply.

VI. DESIGN IMPLEMENTATION

Many of the components and sensors to implement theinstrumentation system weren’t available at local market, sothey were imported.

The PCB16 design was developed with a CAD17 softwarefixing the size to the internal space in the model aircraftand distributing the elements to avoid many vias that couldaffect the voltage signals, especially the signal from thethermocouple that is of the order of “mV”.

An image of the implemented PCB is shown in Fig.7

VII. DESIGN VALIDATION

The validation process has three main parts: Static vali-dation, dynamic validation against controlled processes, andvalidation on the model aircraft.

14Component Object Model15Superficial Mount16Printed Circuit Board17Computer Aided Design

Figure 7. Implemented PCB fixed into the model Aircraft

A. Static Validation

It is necessary to adjust the theoretical transfer functions forthe analogical voltages detected with the ADC module in theµC as the sensors and ICs used for acquire and adequate theflying variables have some variations respect to the standardsgiven in their data-sheets.

These adjustments are done by assuming an offset voltagefor all the transfer function, which is calculated by taking setpoints for the analog variables equal to those present at systemstart. E.g. the output voltage for the differential pressure sensorat 0 ∆Pa is supposed to be 0.6“VDC”, but it is actually0.72“VDC”, so an offset of 0.12“VDC” is assumed to be addedto the transfer function.

B. Dynamic Validation

As an unique reference point is not enough to adjust allthe dynamic range of the sensing modules of this application,some controlled testings were carried out.

1) Thermocouple Test: An Intelligent Temperature SensorFLUKE 52 II and a soldering iron with temperature controllerwere used to analyze the temperature range of this applica-tion. The soldering iron was set at different values while itstemperature was being registered with the FLUKE 52 II andwith the thermocouple of the implemented system.

In this way, some data was extracted and the slope of thetransfer function was adjusted.

2) Speed Test: A simulation of the air flux that will affectthe pitot tube must be done. But this air flux should be equalto the one affecting an airplane in movement, so a motor witha speed controller and a propeller was not a good option asthe air displacement generated by different types of propellersat the same speed is not equal.

In this test, the whole system was mounted in the roof ofa car with digital speedometer. The measured analog voltageand the calculated speed were transmitted to a ground PC andtwo synchronized videos were recorded. The results showedthat speed calculation is fairly good.

3) RPM Meter Test: The software code to measure theperiod of the pulsing signal from the RPM Meter was testedusing a Signal Generator configured to have a square-waveoutput with variable duty cycle. This test had very good resultsand it demonstrated that RPM measuring with this softwarecode is duty-cycle careless.

For the second test, an equivalent engine to the one in themodel aircraft is used. The hall-sensor is attached to the engine

by the back part to detect the counterweight of the crankshaftas in the real implementation. Using a digital tachometerthe measured value was compared to different engine speedsgiving the same good results than in the first test.

4) Altitude Test: The altitude measured with the imple-mented system was compared with a GPS18 receptor, referenceA-1035H made by VincoTech, at different altitudes aroundthe city. With this data, the slope was modified, so theimplemented system could get a better exactitude at altitudemeasurement.

C. Validation on Model Aircraft

The instrumentation system was mounted into the modelaircraft and some test were carried out before taking in intothe air.

First the RF link was established and the engine was turnedon, so it could be checked if there were any interference dueto the mechanical stress of the engine. In this test there wereno problems at all.

Then the engine was accelerated so the RPM value and theengine temperature could be checked.

Finally, some short flights were made to check the changein the speed and in the altitude values.

However, long period flights or long distance flights werenot made because a method to measure the flying distance ofthe model aircraft were not available at the moment.

VIII. ACKNOWLEDGMENT

The authors would like to thank Grupo ERA and to teacherAlfredo Acevedo for his wise advices. Thanks to DonovanPineda, Yuri Mejia, Marlon Velasquez, Carolina Viviescas andAdriana Bernal for their help and support in making thiswork possible. The Universidad Industrial de Santander for itworking labs.

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18Global Position System