In-Flight Lift and Drag Estimation of an Unmanned Propeller-Driven Aircraft

aerospace

Article In-ﬂight Lift and Drag Estimation of an Unmanned Propeller-Driven Aircraft

Dominique Paul Bergmann 1,*, Jan Denzel 1, Ole Pfeiﬂe 2, Stefan Notter 2, Walter Fichter 2 and Andreas Strohmayer 1

1 Institute of Aircraft Design (IFB), University of Stuttgart, Pfaffenwaldring 31, 70569 Stuttgart, Germany; [email protected] (J.D.); [email protected] (A.S.) 2 Institute of Flight Mechanics and Controls (iFR), University of Stuttgart, Pfaffenwaldring 27, 70569 Stuttgart, Germany; ole.pfeiﬂ[email protected] (O.P.); [email protected] (S.N.); walter.ﬁ[email protected] (W.F.) * Correspondence: [email protected]; Tel.: +49-711-685-60342

Abstract: The high-power density and good scaling properties of electric motors enable new propulsion arrangements and aircraft configurations. This results in distributed propulsion systems allowing to make use of aerodynamic interaction effects between individual propellers and the wing of the aircraft, improving flight performance and thus reducing in-flight emissions. In order to systemati- cally analyze these effects, an unmanned research platform was designed and built at the University of Stuttgart. As the aircraft is being used as a testbed for various flight performance studies in the field of distributed electric propulsion, a methodology for precise identification of its performance characteristics is required. One of the main challenges is the determination of the total drag of the aircraft to be able to identify an exact drag and lift polar in flight. For this purpose, an on-board

measurement system was developed which allows for precise determination of the thrust of the aircraft which equals the total aerodynamic drag in steady, horizontal ﬂight. The system has been

Citation: Bergmann, D.P.; Denzel, J.; tested and validated in flight using the unmanned free-flight test platform. The article provides an Pfeifle, O.; Notter, S.; Fichter, W.; overview of the measuring system installed, discusses its functionality and shows results of the flight Strohmayer, A. In-flight Lift and Drag tests carried out. Estimation of an Unmanned Propeller-Driven Aircraft. Aerospace Keywords: unmanned aircraft; thrust determination; flight testing; e-Genius-Mod; free-flight 2021, 8, 43. https://doi.org/10.3390/ wind tunnel aerospace8020043

Academic Editor: Roberto Sabatini Received: 14 December 2020 1. Introduction Accepted: 1 February 2021 While investigating the flight performance of new aircraft configurations or tech- Published: 6 February 2021 nologies, knowledge about aerodynamic and flight mechanical parameters is required to understand their impact on the aircraft. From the perspective of aircraft design, one main Publisher’s Note: MDPI stays neutral focus is the impact on flight performance when investigating for example the effects of with regard to jurisdictional claims in new propulsion technologies or aircraft configurations. To investigate and assess the flight published maps and institutional affil- performance, the determination of lift and drag is of major importance. Especially the iations. determination of drag from flight tests with an unmanned propeller aircraft is a challenging task. Flight tests with unmanned aircraft systems (UAS) have become increasingly important in recent years. Scaled platforms are not only used as payload carriers, but also for the Copyright: © 2021 by the authors. analysis of novel aircraft configurations or for the assessment of unconventional propulsion Licensee MDPI, Basel, Switzerland. systems. This article is an open access article The lower costs as well as the reduction of risks especially in the area of unconven- distributed under the terms and tional configurations militate in favor of using unmanned systems. Typical examples for conditions of the Creative Commons Attribution (CC BY) license (https:// demonstrators used to study flight dynamic effects are the AlbatrossONE [1] of Airbus and creativecommons.org/licenses/by/ the platform of the project FLEXOP (flutter free flight envelope expansion for economical 4.0/). performance improvement) [2]. In the case of the AlbatrossONE, gust loads are minimized

Aerospace 2021, 8, 43. https://doi.org/10.3390/aerospace8020043 https://www.mdpi.com/journal/aerospace Aerospace 2021, 8, 43 2 of 16

by structural interventions in the area of the wing tip, while aeroelastic effects on the wing are observed with the help of the carrier platform FLEXOP. To investigate the impact on flight performance of new technology on an aircraft configuration, a direct comparison between base line configuration and modified aircraft is of interest, as demonstrated by NASA with its Area-I Prototype Technology Evaluation and Research Aircraft (PTERA) [3]. In the research, the PTERA platform was modified to investigate for example combined circulation control [4] and a spanwise adaptive wing [5]. Unmanned aircraft are an important tool to investigate new aircraft configurations and novel aviation technologies at University of Stuttgart, [6]. New technologies or concepts can be flight-tested with little effort, low costs and manageable project risks after initial theoretical investigation. This way, research concepts can reach a much higher feasibility, of particular interest to industry, to close the gap between upstream research and industrial exploitation. This can build a bridge, demonstrating technologies at a higher technology readiness level (TRL). Innovative ideas can be demonstrated and validated in a relevant environment which corresponds to a TRL 5/6. However, there are limitations due to the degree of scaling of the technologies under consideration. A disadvantage for the investigation of flight performance with scaled unmanned aircraft is the limited data available for the different aircraft system components. When investigating the flight performance, propeller characteristics and efficiency of the individual components of the propulsion system are of specific interest. The Institute of Aircraft Design (IFB) at the University of Stuttgart has developed the “e-Genius-Mod” test platform based on this background. The platform is modelled on a scale of 33.3% of the electrically powered “e-Genius” aircraft [7], which is also designed, manufactured and operated by the institute. In this particular case, the UAS was realized as Froude-scaled version [6]. Due to the modular design of the testbed, different configurations can be easily realized, which allows an adaptation to different applications. Furthermore, the geometry of the fuselage offers ideal conditions for accommodating various payloads. For the purpose of identifying the aircraft and measuring flight performance, it is essential to map the aerodynamic parameters of the system. The most important parameters are the aircraft drag and the lift polar. From the correlation of the coefficients for lift and drag of the aircraft, a direct statement about the performance and energy consumption can be derived. The gliding characteristics, the required power installed and the potential range are important parameters that are required to compare different propulsion concepts and identify their impact for a prospective aircraft design. In order to obtain reliable drag and lift polar, manned test aircraft often would carry out comparative flights with calibrated systems. Especially in the field of gliding, where a precisely identified system is very important, such survey flights are still state of the art today. In the unmanned area, such survey flights are difficult to carry out. For UAS, a method for a glider model is described by Edwards [8]. Due to the induced drag of the propeller system, this method is not useful for the proposed identification of propeller-driven UAS. A method to identify drag of a propeller aircraft was developed by Norris and Bauer [9]. Alternatively, if available, a model for the propeller wind-milling drag or thrust could be used to apply this method. For the e-Genius-Mod testbed, a different approach was taken. In order to record the performance data, a measurement system was developed that is capable of recording the thrust values required during the flight. Such systems have not yet been installed in large UAS, but have only been used sporadically on manned aircraft. For the aerodynamic characterization of smaller UAS, on-board thrust measurements have been performed in the past, such as presented by M. Bronz and G. Hattenberger [10]. However, for the identification of the flight performance and to show the feasibility of novel technologies in aviation the measurement of drag is essential. Although the thrust values for the operation of a propeller-driven system can be measured without any problems when stationary, the thrust values during flight are not Aerospace 2021, 8, x FOR PEER REVIEW 3 of 16

Aerospace 2021, 8, 43 Although the thrust values for the operation of a propeller-driven system can be 3 of 16 measured without any problems when stationary, the thrust values during flight are not comparable due to the induced flow to the propeller without prior identification of the propeller via wind tunnel experiments. The efficiency of the drive also changes signifi- comparable due to the induced flow to the propeller without prior identification of the cantly under different flow conditions. Therefore, a direct, mechanical thrust measure- propeller via wind tunnel experiments. The efficiency of the drive also changes significantly ment in flight is of significant advantage. under different flow conditions. Therefore, a direct, mechanical thrust measurement in flight is of significant advantage. 1.1. Challenge of Current Research Using Unmanned Aircraft as Free-Flight Test Platform A current 1.1.research Challenge topic of at Current the University Research Usingof Stuttgart Unmanned is how Aircraft to use as unmanned Free-Flight pro- Test Platform peller-driven aircraftA currentas a test research platform topic to determine at the University their flight of Stuttgartperformance is how without to use unmanned knowing specificpropeller-driven data of propeller aircraft and powertra as a test platformin. This ability to determine is for example their flight required performance to without compare differentknowing propulsion specific configurations data of propeller like andWing powertrain. Tip Propellers This (WTP) ability to is fora basic example required configuration into flight. compare Based different on the propulsion collected configurationsflight data, the like estimated Wing Tip benefits Propellers of this (WTP) to a basic technology in flightconfiguration should be in validated flight. Based and demonstrated. on the collected For flight this purpose, data, the the estimated basic in- benefits of this flight measurementtechnology system in of flight the test should platform be validated is extended and with demonstrated. a system to Formeasure this purpose, the the basic thrust of the propellerin-flight in-flight. measurement system of the test platform is extended with a system to measure For this, thethe assumption thrust of the was propeller made in-flight.that in steady horizontal flight the thrust performed by the propellerFor this, corresponds the assumption to the wasaerodynamic made that drag in steady acting horizontal on the aircraft. flight the thrust performed by the propeller corresponds to the aerodynamic drag acting on the aircraft. 1.2. The Modular Test Platform e-Genius-Mod The free-flight1.2. The platform Modular e-Genius-Mod Test Platform e-Genius-Mod(Figure 1) is established as a technology test bed to demonstrate Thenew free-flight technologi platformes for future e-Genius-Mod aircraft design (Figure in1 a) isrelevant established environment. as a technology test bed The UAS test platformto demonstrate [6] is used new for technologies academic and for futureinnovative aircraft research design projects in a relevant to inves- environment. The tigate scaling similaritiesUAS test platform of free flight [6] is models used for and academic to demonstrate and innovative new technologies research projects up to to investigate TRL 6. The modularscaling design similarities of the of test free bed flight is modelsideally suited and to for demonstrate the investigation new technologies of new up to TRL 6. aircraft configurationThe modular solutions design for distributed of the test bedelectric is ideally propulsi suitedon systems. for the investigationThe size of the of new aircraft aircraft with a configurationwingspan of 5.62 solutions m, maximum for distributed take-off electric weight propulsion of 40 kg and systems. a payload The sizeca- of the aircraft pacity of morewith than a 10 wingspan kg [6] is ofsuited 5.62 m,to perform maximum prospective take-off weight investigations. of 40 kg and The a maxi- payload capacity of mum possible moreflight thantime 10 is kgup [ 6to] is100 suited min, to depending perform prospective on the particular investigations. battery Thecapacity maximum possible and the payloadflight weight. time An is up overview to 100 min, of the depending technical on data the of particular the e-Genius-Mod battery capacity is sum- and the payload marized in Tableweight. 1. An overview of the technical data of the e-Genius-Mod is summarized in Table1.

Figure 1. Free-flight test Figureplatform 1. e-Genius-Mod.Free-ﬂight test platform e-Genius-Mod.

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Table 1. Technical data of the e-Genius-Mod [6].

Aircraft Parameter Value Wing span 5.62 m Wing area 1.56 m2 Aspect ratio 20.2 Length 2.95 m Take of mass * 34.9 kg Electric drive power 5 kW (Plettenberg Terminator 30/8 Evo) Max. thrust 156 N Propeller RASA 24 × 12 Design speed 24.8 m/s Battery capacity * 42 Ah (44.4 V) * ﬂight test characteristic.

The test bed e-Genius-Mod is an extension of the full-scale aircraft for further investigation of electric flight and new aviation technologies. For this reason, the test platform is equipped as a free flight wind tunnel. In order to perform measurements in steady, horizontal flight, an autopilot is used to steer the aircraft along a predefined path of constant altitude and velocity, as described in [11]. The measurement system ensures the synchronous logging of all relevant variables (Table2), as further detailed in Section3.

Table 2. Basic measurement equipment of the free-ﬂight wind-tunnel.

Variable Sensor Type Unit linear accelerations IMU rotation rates Board computer IMU (Pixhawk4) position (IMU, GPS, magnetometer, attitude barometer) estimation via velocity Kalman filter true air speed AirDataBoom angle of attack 5 hole-probe (Vectoflow/VectoDAQ angle of sideslip temperature sensor Air Data Computer) air density deflection angle of the control magnet sensor Actuator (Volz surfaces rotor sensor DA15N)

The basic free flight measurement equipment can be easily expanded by connecting additional sensors to the dedicated bus system used solely for measurement data. Electric propulsion systems allow for new design alternatives in terms of propulsion integration and aircraft design. To investigate and demonstrate new concepts, a modular testbed like the unmanned scale model is a very useful and flexible tool. The demonstration of innovative concepts like distributed propulsion on a manned aircraft would be expensive and time intensive and not useful for research with open-ended findings at this early stage. With its modular airframe design, the e-Genius-Mod is the basis for an efficient and systematic research of the various effects of distributed propulsion. Onboard thrust measurements on a UAS pose a particular challenge. On the one hand, it must be possible to carry out reliable, calibrated measurements, and on the other hand, the sensor systems are subject to narrow limits in terms of dimensions and weight. For this purpose, sensor systems specifically designed for the e-Genius Mod were developed, calibrated and tested in wind tunnel experiments before installation. The thrust values are in many respects informative for the evaluation of the flight controller itself, as well as for the assessment of the efficiency of an engine. In general, they establish a direct relationship to aerodynamic quality and, in case the electrical power consumption is known, allow a direct statement about the overall efficiency of the Aerospace 2021, 8, 43 5 of 16

corresponding powertrain. Especially when considering several distributed engines, it is possible to make a reliable statement about the overall energy balance onboard the platform. The aim of the thrust measurements is to prove the expected positive effects of the distributed engines in terms of quality, and, in further steps, to make statements about where on the aircraft they can be used most efﬁciently. Furthermore, previously performed simulations are to be validated with these measured values.

2. Research Objective To assess and validate the impact of a novel configuration or technology in flight, knowledge of the flight performance is of vital importance. Required basic information in aircraft design are the aerodynamic coefficients for lift (CL) and drag (CD) of an aircraft with respect to the angle of attack (AoA). Therefore, the objective of this research is the identification of the CL − CD, CL − AoA and CD − AoA polars in-flight. For the investigation of these coefficients in-flight the following well-known approach is proposed as a starting point. While the investigations are carried out under cruise conditions, it is assumed as a basis for the investigations that the measurements are carried out in a non-accelerated (steady) horizontal flight. This allows to establish the balance of forces in flight direction and perpendicular to it respectively:

DRAG (D) = THRUST (T) (1)

LIFT (L) = WEIGHT (W) (2) Thrust will be directly measured between engine and engine mount. The measured force corresponds to the force acting on the aircraft caused by the propeller thrust. With ρ T ≈ D = ∗ v2 ∗ C ∗ S (3) 2 D and the knowledge about the true air speed (v) and air density (ρ), we can directly determine the drag coefficient (CD) of the aircraft by assuming drag (D) and thrust (T) as balanced. S describes the reference wing area of the test platform. True air speed and air density will be measured with the air data boom installed in the nose of the aircraft (Table2) . A small deviation has to be considered by the frictional forces of linear bearings in the thrust measurement unit which can be eliminated by a calibration. The lift coefficient (CL) can be determined by the following equation: ρ W = L = ∗ v2 ∗ C ∗ S (4) 2 L Since the aircraft is electrically powered by a battery system, the weight is constant throughout the flight. To identify the corresponding angle of attack, the air data boom is used. Angle of attack and angle of sideslip are measured with a five-hole probe. As the measured values of the different sensors are synchronised, the coefficients of lift and drag can be described directly in relation to the AoA. The measured AoA has only to be corrected by its installation position in relation to the wing. In this way, the direct connection between AoA and the corresponding lift and drag values should be representative in flight. Larger values of lift and drag are to be expected with an increasing of the AoA.

3. Approach for the Flight Test Scenario The approach for the flight test scenario is based on the assumptions made in Section2 to investigate the flight performance under the condition of a steady horizontal flight and with a zero-wind condition for the atmosphere. However, considering the reality of free-flight tests, perfect atmospheric conditions will never be achieved. To meet these requirements to some degree, the flight tests are carried out in the early morning on days with calm atmosphere. The impact of atmospheric Aerospace 2021, 8, x FOR PEER REVIEW 6 of 16

Aerospace 2021, 8, 43 However, considering the reality of free-flight tests, perfect atmospheric conditions6 of 16 will never be achieved. To meet these requirements to some degree, the flight tests are carried out in the early morning on days with calm atmosphere. The impact of atmos- phericdisturbances disturbances was estimated was estimated in [12]. Toin realize[12]. To a statisticalrealize a accuracystatistical of accuracy the data measuredof the data in measuredflight, the in data flight, for athe single data measurementfor a single me pointasurement is collected point overis collected a minimum over a of minimum four legs ofwith four two legs different with two flight different directions. flight directions. ToTo determine determine the the thrust thrust during during flight, flight, a a measuring system system is is installed installed between between the the electricelectric engine engine and and the the engine engine mount mount to to measure measure the the tensile tensile forces. forces. The The occurring occurring tensile tensile forcesforces correspond correspond to to the the thrust thrust generated generated by by the the propulsion propulsion system. system. TheThe thrust thrust measurement measurement is is particularly particularly releva relevantnt due due to to its its direct direct correlation correlation with with the the aerodynamicaerodynamic drag. drag. The The investigations investigations are are carri carrieded out out under under cruise cruise cond conditions,itions, i.e., i.e., steady, steady, horizontalhorizontal flight. flight. For For the the investigation, investigation, a a flig flightht track track is is chosen chosen which which allows allows the the longest longest possiblepossible horizontal legs.legs. TheThe limitinglimiting factor factor for for the the tests tests is theis the current current regulatory regulatory framework, framework,which which specifies specifies a maximum a maximum flight envelopeflight envelope for the for tests the in tests the in permitted the permitted airspace. airspace. TheThe measurement flightsflights are are performed performed with with an autopilotan autopilot in control in control to achieve to achieve steady, steady,horizontal horizontal flight and flight high and repeatability. high repeatability. For the testFor the flight test a flight circuit awith circuit maximized with maximized straight straightsegments segments (legs) in-between (legs) in-between is chosen. is Givenchosen the. Given airspace the restrictions,airspace restrictions, the leg on the which leg theon whichmeasurements the measurements can be made can isbe nearly made 1000is nearly m long. 1000 Figure m long.2 represents Figure 2 represents the flight the path flight of a pathmeasurement of a measurement flight at 300flight m at altitude. 300 m altitude.

FigureFigure 2. 2. FlightFlight track—sections track—sections of of measurement measurement ( (greengreen);); level level off off sections sections ( (orangeorange).).

TheThe legs legs are are divided divided into into two two segments. segments. The The first first segment segment is is the the “level “level off off section”. section”. AfterAfter the the turn, turn, airspeed, airspeed, altitude altitude and and attitude attitude are are stabilized stabilized in in this section. Data Data in in this this sectionsection will will not not be be considered. considered. In In the the sectio sectionn of of measurement, measurement, a a stabilized stabilized flight flight attitude attitude isis expected expected and and the the data data will will be be used used for for the the an analysisalysis of of lift lift and and drag. drag. As As there there will will be be no no perfectperfect steady steady horizontal flightflight withwith constantconstant altitude altitude and and airspeed, airspeed, limits limits for for the the deviation devia- tionin roll in roll angle, angle, airspeed airspeed and and altitude altitude are are set, set, to assign to assign a confidence a confidence rating rating to the to the single single leg legthat that is considered is considered in thein the evaluation. evaluation. To optimizeTo optimize the the steady steady horizontal horizontal flight, flight, the the autopilot auto- pilotcontrols controls a constant a constant velocity velocity and altitude.and altitude. 4. Measurement Unit The thrust measurement unit consists of two parts. The sensor unit which contains the sensor itself for direct measurement of the thrust and a self-developed data acquisition Aerospace 2021, 8, x FOR PEER REVIEW 7 of 16

4. Measurement Unit The thrust measurement unit consists of two parts. The sensor unit which contains the sensor itself for direct measurement of the thrust and a self-developed data acquisition unit that contains the amplifier module and the interfaces to connect further sensors. The Aerospacedata2021, 8acquisition, 43 unit is also the interface to the data storage. 7 of 16 Recorded values at the propulsion system during flight: • Thrust • RPM of the propellerunit that contains the amplifier module and the interfaces to connect further sensors. The • data acquisition unit is also the interface to the data storage. Engine current Recorded values at the propulsion system during flight: • Engine voltage• Thrust Besides collecting• theRPM data of the with propeller high precision, the complete measurement system is required to feature •a lightweightEngine current design for use on the UAS. This ensures that the use of • Engine voltage the testbed for further studies is not limited by the weight of the measurement unit. Besides collecting the data with high precision, the complete measurement system is required to feature a lightweight design for use on the UAS. This ensures that the use of 4.1. Sensor Unit the testbed for further studies is not limited by the weight of the measurement unit. The sensor unit is based on a standard tension/pressure sensor (HBM U9C 100N). For 4.1. Sensor Unit the sensor unit to be used on the unmanned test platform, the main focus is on a robust The sensor unit is based on a standard tension/pressure sensor (HBM U9C 100N). For and compact lightweightthe sensor design. unit to The be used unit on is the designed unmanned to test allow platform, for the mainsensor focus to isbe on in- a robust stalled without torqueand and compact bending lightweight moments design. introduced The unit is designed by the toengine. allow for the sensor to be installed An engine connectionwithout torque unit andis being bending developed moments introducedfor this purpose. by the engine. This consists of an engine connection plate,An a engineforce introduction connection unit bolt is being and developedlinear guide for this pins purpose. to absorb This torsion consists of an and bending loads. engineThese connection loads are plate, directed a force from introduction the linear bolt andguiding linear pins guide into pins tothe absorb linear torsion and bending loads. These loads are directed from the linear guiding pins into the linear bearings installed inbearings the load installed transmission in the load frame. transmission For the frame. inflight For the thrust inflight measurement, thrust measurement, the the sensor unit is installedsensor directly unit is installedbetween directly the engine between mount the engine and mount the engine and the (Figure engine (Figure 3). 3).

FigureFigure 3. Integrated 3. Integrated thrust thrust measurementmeasurement sensor sensor unit; unit; 1—Sensor 1—Sensor Unit; 2—Engine Unit; 2—Engine Mount; 3—Electric Mount; Engine. 3—Elec- tric Engine. Due to its design, the sensor unit allows two separate load paths (Figure4). Loads will be introduced in the sensor unit via the engine connection plate. The plate is linearly Due to its design,mounted the sensor and transmits unit allows the force two induced separate by the propellerload paths thrust (Figure as tension 4). forceLoads directly will be introduced into the the sensor.sensor Theunit torque via the and engine the bending connection loads resulting plate. The from plate the angular is linearly and linear mounted and transmitsacceleration the force and induced vibrations by of the the propeller engine are thrust derived as via tension the linear force bearings directly into the to the sensor. The torquestructure and of the the sensor bending unit (housingloads resulting and frames). from This the way angular the thrust and can linear directly be derived from the measured tensile force. acceleration and vibrations of the engine are derived via the linear bearings into the struc- ture of the sensor unit (housing and frames). This way the thrust can directly be derived from the measured tensile force. Aerospace 2021, 8, x FOR PEER REVIEW 8 of 16

tension rod (traction stop) friction bearing tension sensor AerospaceAerospace 20212021,, 88,, x 43 FOR PEER REVIEW 8 8 of of 16 16 connecting frame housing tension rodforce (traction introduction stop) bolt friction bearingengine connection plate tension sensorload transmission frame connecting linearframe guide pins housing linear bearing force introduction bolt engine connection plate load transmission frame linear guide pins linear bearing

Figure 4. Sensor Unit (sectional view).

4.2. Data Acquisition Unit Figure 4. Sensor Unit (sectional view). FigureThe 4. Sensor task Unit of (sectionalthe data view). acquisition unit (Figure 5) is the acquisition and processing of 4.2. Data Acquisition Unit 4.2.data Data from Acquisition the propulsion Unit system and the transmission of the data to the board computer for synchronousThe task of the logging. data acquisition An STM32 unit (Figure microprocess5) is the acquisitionor is used and to processingconvert analog of signals, to dataThe from task the of propulsion the data acquisition system and unit the (Figure transmission 5) is the of theacquisition data to the and board processing computer of dataencodefor synchronousfrom and the propulsiontime logging. stamp system An the STM32 data and the microprocessorand tran tosmission handle is oftransmission used the data to convert to the to analogboard the cecomputer signals,ntral board to computer forviaencode synchronous an RS485 and time bus logging. stamp (Figure the An data STM32 6). and While microprocess to handle this transmissionis sufficientor is used toto to theconvert handle central analog boardthe sensorsignals, computer toinput from mag- encodeneticvia an rpm RS485and timemeasurements bus stamp (Figure the6). data While and and this currentto is handle sufficient andtransmission to voltag handlee theto sensors, the sensor central input the board fromforce computer magnetic sensor outputs need via an RS485 bus (Figure 6). While this is sufficient to handle the sensor input from mag- torpm be measurements amplified first. and currentThis is and done voltage using sensors, a miniaturized the force sensor amplifier outputs need module to be (GSV-6CPU) netic rpm measurements and current and voltage sensors, the force sensor outputs need installedamplified on first. the This Data is done Acquisition using a miniaturized Unit. Additional amplifier moduleinterfaces (GSV-6CPU) on the data installed acquisition board toon be the amplified Data Acquisition first. This Unit. is done Additional using a interfaces miniaturized on the amplifier data acquisition module board(GSV-6CPU) allow to installedallowconnect to on further connect the Data sensors furtherAcquisition that mightsensors Unit. be Additional requiredthat might by interfaces future be required studies on the using databy futureacquisition the e-Genius-Mod studies board using the e-Ge- allownius-Modtest platform. to connect test With furtherplatform. a weight sensors ofWith only that a 28might weight g, the be data requiredof only acquisition by28 futureg, unitthe studies (boarddata acquisition withoutusing the cable) e-Ge- unit is (board with- nius-Modoutperfectly cable) suitedtest is platform. perfectly for use With in suited UAS a weight without for useof restrictingonly in 28UAS g, thethe without payloaddata acquisition restricting capacity. unit the(board payload with- capacity. out cable) is perfectly suited for use in UAS without restricting the payload capacity.

Figure 5. Data Acquisition Unit.

Figure 5. Data Acquisition Unit. Figure 5. Data Acquisition Unit. Figure 6. Information flow of the Data Acquisition Unit.

Figure 6. Information flow of the Data Acquisition Unit. Aerospace 2021, 8, x FOR PEER REVIEW 8 of 16

tension rod (traction stop) friction bearing tension sensor connecting frame housing force introduction bolt engine connection plate load transmission frame linear guide pins linear bearing

Figure 4. Sensor Unit (sectional view).

4.2. Data Acquisition Unit The task of the data acquisition unit (Figure 5) is the acquisition and processing of data from the propulsion system and the transmission of the data to the board computer for synchronous logging. An STM32 microprocessor is used to convert analog signals, to encode and time stamp the data and to handle transmission to the central board computer via an RS485 bus (Figure 6). While this is sufficient to handle the sensor input from magnetic rpm measurements and current and voltage sensors, the force sensor outputs need to be amplified first. This is done using a miniaturized amplifier module (GSV-6CPU) installed on the Data Acquisition Unit. Additional interfaces on the data acquisition board allow to connect further sensors that might be required by future studies using the e-Ge- nius-Mod test platform. With a weight of only 28 g, the data acquisition unit (board without cable) is perfectly suited for use in UAS without restricting the payload capacity.

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Figure 5. Data Acquisition Unit.

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Figure 6. InformationFigure flow of 6. theInformation Data Acquisition ﬂow of Unit. the Data Acquisition Unit.

5. Test Flight Analysis 5. Test Flight Analysis In order to obtain repeatable measurement flights, the flights are automatically In order to obtain repeatable measurement flights, the flights are automatically guided guided along a previously planned flight path. The autopilot system [11] adjusts the flight along a previously planned flight path. The autopilot system [11] adjusts the flight condi- condition to different, constant flight speeds at an altitude above mean sea level of 650 m. tion to different, constant flight speeds at an altitude above mean sea level of 650 m. This This results in stable conditions in stationary horizontal flight, during which the corre- results in stable conditions in stationary horizontal flight, during which the corresponding sponding input variables of the drive system are recorded. The available endurance of input variables of the drive system are recorded. The available endurance of 85 min for one 85 min for one test flight allows for more than 68 measuring sections, which are seg- test flight allows for more than 68 measuring sections, which are segmented into phases mented into phases with different flight velocities. With this, all measurements required with different flight velocities. With this, all measurements required for a characterization for a characterization of the aerodynamic parameters of a configuration can be carried out of the aerodynamic parameters of a configuration can be carried out in one single flight inFigure one single7 provides flight a Figure synopsis 7 provides of the legs a synops and Tableis of3 thepresents legs and the Table commanded 3 presents velocities the com- to mandedidentify velocities the coefficients to identify for lift the and coeffi dragcients in flight. for lift and drag in flight.

Figure 7. Section of the flight path from a single measurement flight. Figure 7. Section of the flight path from a single measurement flight. Table 3. Mean standard deviation of velocity and altitude. Table 3. Mean standard deviation of velocity and altitude. Velocity (m/s) Altitude (m) Velocity (m/s) Altitude (m) commandedCommanded standard Standard deviation Deviation commanded Commanded standard Standard deviation Deviation 2020 0.23 0.23 0.090.09 2121 0.23 0.23 0.11 0.11 22 0.21 0.12 2224 0.21 0.31 0.12 0.12 300 2426 0.31 0.25 0.12 0.13 300 2628 0.25 0.22 0.13 0.11 2830 0.22 0.22 0.11 0.08 32 0.20 0.08 30 0.22 0.08 32 0.20 0.08 Analysis of Static Horizontal Flight AnalysisIn orderof Static for Horizontal the evaluation Flight to be considered valid using the mean value method, the “section of measurement” must be carried out at a velocity and an altitude as constant as In order for the evaluation to be considered valid using the mean value method, the possible. The evaluation of the individual legs shows that both velocity and altitude are “section of measurement” must be carried out at a velocity and an altitude as constant as measured with only a very small standard deviation (Table3) depending on the flight speed. possible. The evaluation of the individual legs shows that both velocity and altitude are As expected, high fluctuations were detected in the recorded propeller speed, which measured with only a very small standard deviation (Table 3) depending on the flight is also reflected in the measured thrust values. This is a result of controlling the airspeed speed. by using the thrust. The resulting control action can be seen in the resulting fluctuations As expected, high fluctuations were detected in the recorded propeller speed, which is also reflected in the measured thrust values. This is a result of controlling the airspeed by using the thrust. The resulting control action can be seen in the resulting fluctuations in measured thrust values (Figure 8). They do not mainly represent a scattering in the Aerospace 2021, 8, 43 10 of 16

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in measured thrust values (Figure8). They do not mainly represent a scattering in the aerodynamicaerodynamic drag,drag, butbut moreovermoreover the the necessary necessary control control action action to keepto keep the the airspeed airspeed constant. constant.This is This particularly is particularly the case the during case during the two the turns. two turns.

FigureFigure 8. Example:Example: Basic Basic measurements measurements on on one one single single leg leg (with (with TAS TAS 28 m/s). The standard deviations caused by this must be taken into account accordingly in the The standard deviations caused by this must be taken into account accordingly in the evaluation and interpretation of the measured values. Nevertheless, the measured values evaluation and interpretation of the measured values. Nevertheless, the measured values on average give a good picture of the expected results and provide an excellent indication on average give a good picture of the expected results and provide an excellent indication of the performance and aerodynamic quality of the aircraft. of the performance and aerodynamic quality of the aircraft. For a first estimation and interpretation of the flight results, the values for thrust, angle For a first estimation and interpretation of the flight results, the values for thrust, of attack, lift and power are approximated as a function of 2nd order in dependence of the angle of attack, lift and power are approximated as a function of 2nd order in dependence velocity as expected from the theoretical consideration. The assumption is based on the of the velocity as expected from the theoretical consideration. The assumption is based on approximate relationship, that the lift increases linearly in relation to the AoA and the drag the approximate relationship, that the lift increases linearly in relation to the AoA and the increases quadratic to the lift. Even if this procedure involves a degree of uncertainty in the drag increases quadratic to the lift. Even if this procedure involves a degree of uncertainty interpretation of the results, it allows for an initial analysis of the aircraft characteristics, to in the interpretation of the results, it allows for an initial analysis of the aircraft character- provide input data for the aircraft design process. istics, to provide input data for the aircraft design process. 6. Results 6. Results This section presents the results of the test flight presented in the previous section. ConsideringThis section that thepresents UAS shouldthe results be used of the as atest tool flight in aircraft presented design, in the the analysis previous focuses section. on Consideringthe most interesting that the aspects UAS should for a design be used engineer. as a tool Figures in aircraft9 and design, 10 show the the analysis coefficients focuses of onlift the and most drag interesting obtained from aspects the measurement for a design engineer. flights and Figures the approximated 9 and 10 show resulting the coeffi- polar cientscurve. of As lift expected, and drag due obtained to the rpmfrom variationthe measurement described flights above, and a significantly the approximated increased re- sultingvariation polar of the curve. identified As expected, drag coefficient due to the can rpm be seenvariation in Figure described 10. Nevertheless, above, a significantly the values increasedcan be used variation for first of design the identified calculations. drag coefficient can be seen in Figure 10. Neverthe- less, the values can be used for first design calculations. Aerospace 2021, 8, x FOR PEER REVIEW 11 of 16

Aerospace 2021, 8, x FOR PEER REVIEW 11 of 16 Aerospace 2021, 8, 43 11 of 16

Figure 9. Lift polar. FigureFigure 9. Lift 9. Lift polar. polar.

FigureFigure 10. 10.DragDrag polar. polar. Figure 10. Drag polar. EvenEven if ifonly only a acertain certain regime regime of of the dragdrag polarpolar andand the the thrust thrust to to velocity velocity curve curve is avail- is available,able,Even design designif only points pointsa certain of of interest interestregime (Table (Tableof the4) can drag4) can be polar estimatedbe estimated and the with withthrust the the available to availablevelocity identiﬁcation. curveidentifi- is available,cation.The pointsThe design points of maximum points of maximu of interest rangem range and (Table and maximum maximum4) can endurancebe estimated endurance can with directlycan thedirectly available be identiﬁed. be identified. identification. The points of maximum range and maximum endurance can directly be identified. TableTable 4. Estimated 4. Estimated points points of max. of max. range range and and max. max. endurance. endurance. Table 4. Estimated points of max. range and max. endurance. Max.Max. Range Range Max. Max. Endurance Endurance Max. Range Max. Endurance VelocityVelocity 20.96 20.96 m/s m/s 20.15 20.15 m/s m/s GlideVelocityGlide Ratio Ratio 20.96 17.96 m/s 17.96 20.15 17.88 17.88m/s Glide RatioCL 17.960.91 17.88 0.98 𝐶 0.91 0.98 C 0.0507 0.0548 𝐶 D 0.91 0.98 Thrust 0.0507 19.06 N0.0548 19.04 N Thrus𝐶 t 19.060.0507 N 19.040.0548 N Thrust 19.06 N 19.04 N Aerospace 2021, 8, x FOR PEER REVIEW 12 of 16

Aerospace 2021, 8, x FOR PEER REVIEW 12 of 16 Aerospace 2021, 8, 43 12 of 16

The point of maximum range is given by the maximum of the glide ratio (𝐶/𝐶) and can be Thedetermined point of directlymaximum from range Figures is given 10 and by 11.the Themaximum point of of maximum the glide ratioendurance (𝐶/𝐶 can) and becan determined be Thedetermined point as point of directly maximum of minimum from range Figures required is given 10 and bythrust the 11. maximum(minThe point drag) of in themaximum Figure glide 12. ratio endurance (CL/CD can) and becan determined be determined as point directly of minimum from Figures required 10 and thrust 11. (min The point drag) of in maximum Figure 12. endurance can be determined as point of minimum required thrust (min drag) in Figure 12.

Figure 11. Glide ratio (CL/CD) to velocity. FigureFigure 11. 11. GlideGlide ratio ratio (C (LC/CL/D)C toD) velocity. to velocity.

FigureFigure 12. Thrust 12. Thrust to velocity. to velocity. Figure 12. Thrust to velocity. With the knowledge of the electrical power and the power output of the propeller in terms of thrust, the overall efﬁciency of the test platform can be determined (Figure 13). With the knowledge of the electrical power and the power output of the propeller in terms ofWith thrust, the knowledgethe overall efficiencyof the electrical of the powertest platform and the can power be determined output of the (Figure propeller 13). in terms of thrust, the overall efficiency of the test platform can be determined (Figure 13). Aerospace 2021, 8, x FOR PEER REVIEW 13 of 16 Aerospace 2021, 8, x FOR PEER REVIEW 13 of 16

Aerospace 2021, 8, 43 13 of 16

Figure 13. Performance and Efficiency of the free-flight test platform for different velocities. FigureFigure 13. 13. PerformancePerformance and and Efficiency Efﬁciency of ofthe the free-flight free-ﬂight test test platform platform for for different different velocities. velocities.

7.7. 7.DiscussionDiscussion Discussion TheTheThe measurementsmeasurements measurements obtainedobtained obtained fromfrom from thethe the flightflight flight teststests tests providedprovided provided thethe the expectedexpected expected resultsresults results withwith with regard to the thrust curves and the resulting polar. As a first comparison between flight test regardregard toto thethe thrustthrust curvescurves andand thethe resultingresulting polar.polar. AsAs aa firstfirst comparisoncomparison betweenbetween flightflight results and a simulation of the wing airfoil, the gradient of lift coefficient versus the angle testtest resultsresults andand aa simulationsimulation ofof thethe wingwing airfoiairfoil,l, thethe gradientgradient ofof liftlift coefficientcoefficient versusversus thethe of attack was assessed. A comparison with the lift of the airfoil simulated in XFoil [13] angleangle ofof attackattack waswas assessed.assessed. AA comparisoncomparison withwith thethe liftlift ofof thethe airfoilairfoil simulatedsimulated inin XFoilXFoil shows a very good agreement with reality (Figure 14). The slight deviation of the slope of [13][13] showsshows aa veryvery goodgood agreementagreement withwith realitrealityy (Figure(Figure 14).14). TheThe slightslight deviationdeviation ofof thethe the two curves results from the influence of the lift distribution on the real wing caused by slopeslope ofof thethe twotwo curvescurves resultsresults fromfrom thethe influencinfluencee ofof thethe liftlift distributiondistribution onon thethe realreal wingwing the boundary vortices. causedcaused byby thethe boundaryboundary vortices.vortices.

FigureFigure 14. 14. ComparisonComparison of ofXFoil XFoil airfoil airfoil simulation, simulation, semi semi empirical empirical method method and and flight ﬂight test test data. data. Figure 14. Comparison of XFoil airfoil simulation, semi empirical method and flight test data.

The amount of the deviation depends on the aspect ratio (AR) of the wing and can be calculated approximately [14] as follows with b describing the wing span and S the reference wing area:

dC AR b2 L = ∗ 2π with AR = (5) dAoA (2 + AR) S Aerospace 2021, 8, 43 14 of 16

The angle of attack for zero lift is found as −3.8◦ from interpolation of the measured values from real ﬂight, which gives a good agreement with the value expected from the simulation. The corresponding installation position of the air data probe in relation to the wing’s angle of attack is thus correctly recorded. An additional validation method was performed, calculating the wing’s lift curve slope with a semi-empirical formula [15], applicable for subsonic design. 2πAR Sexp = CLAoA s F (6) 2 S 2 2 (tan ARmax ) + + AR β + t 2 4 η2 1 β2

Within formula (6) η describes the so-called efficiency of the airfoil and is approximately 0.95 according to [15]. β2 considers the flight velocity respectively the Mach number (M). β2 = 1 − M2 (7) The wing geometry and the influence of the wing are considered in Equation (7) with

ARmaxt , Sexp and F: ARmaxt is the aspect ratio of the wing section with the thickest airfoil chord location. The wing area exposed by the fuselage is considered with Sexp and the lift generated by the wing is approximated with F estimated via the diameter d of the fuselage tube under the wing. According to Raymer [15], these can be calculated as follows.

d 2 F = 1.07 1 + (8) b

To consider also the effects of the winglet the effective aspect ratio of the wing is estimated according to [14] taking into account the height of the winglets h in relation to the wing span. h AR = AR 1 + 1.9 (9) e f f b

The polar of the calculated lift curve slope (CLAoA = 0.0764) is plotted in Figure 14 and the zero crossing is shifted by the estimated = 0.33. In a direct comparison of the part of the lift polar which can be considered as nearly linear between the semi-empirical and the = ( ◦ < < ◦) measured lift curve slope with CLAoA measured 0.0787 for 0 AOA 5 a good match is achieved with only 3.01% deviation of the curve slope. In the simplified approach used in this paper, the possible error in the results must also be considered. A relatively small error can be assumed for the lift coefficient. This error is made up of the accuracy of the air data probe (measurement error flow angle: <1.0◦, velocities: <1.0 m/s) and a possible installation error which is corrected in the post processing. The drag measurement must be viewed with greater caution. The accuracy of the measuring system has been demonstrated under laboratory conditions, the variation of the measured thrust values due to the fluctuating speed controller (see Section5) leads to a high standard deviation (mean 26.56%). Nevertheless, the thrust measuring system generated usable data, which, however, are subject to fluctuations due to controlling thrust values. This must be considered in the evaluation. For further flight tests, control authority for the airspeed control should be reduced in order to ensure that the measured thrust values are smoother and there is less scattering. The proper functioning of the measuring system was verified in laboratory tests before and after the flight tests. The measurement of the thrust finally gives information about the total drag, which is necessary to establish the drag polar and aerodynamic parameters of interest. From the illustrations in the “results” section the corresponding statements regarding the aerodynamic performance of the system can be derived. The values are within the predicted range. As expected, the scaling of the aircraft brings the optima for best range and flight duration very close together. By measuring the thrust during flight while simultaneously Aerospace 2021, 8, 43 15 of 16

observing the electrical power consumption, a direct statement can also be made about the overall efﬁciency of the drive train. It can be seen from Figure 13 that the optimum efﬁciency is found in the range of the speed at which the system is operated in the glide ratio of best range. This is an important step to assess the impact of future design changes to the e-Genius Mod.

8. Conclusions The measurement flights with the unmanned e-Genius-Mod test platform have shown that on-board thrust measurement provide a direct method to characterize the aircraft. The expectations to draw direct conclusions about the aerodynamic performance could be met. With the presented novel measurement system for direct, in-flight thrust measurements, the unmanned platform represents a useful tool in aircraft design as a “flying wind tunnel”. It allows an identification of the flight performance and allows to evaluate the effects of new technologies and configurations. This is of particular interest for new configurations that use distributed electric propulsion to increase efficiency and to enable a future cleaner and greener aviation. To improve the quality of the flight test results in the future and to allow for drag estimation from non-steady flight conditions, the aircraft dynamics will be included in the drag estimation through acceleration measurements from an inertial measurement unit. Using the thrust measurement system, the characterization of the aircraft could be completed to such an extent that further flight tests with modified settings and distributed propulsion systems can follow. Due to the performance measurements on the baseline configuration it will be possible in the future to draw direct comparisons to flight tests with modified configurations. The measurement system used has been verified and can now also be used for further configurations utilizing distributed propulsion. For this, the sensor unit will be miniaturized to be able to measure the thrust of each propulsion unit in flight. The additional measuring technology installed to record the aerodynamically relevant input variables perfectly complements the thrust measurement of the main propulsor. Therefore, the basic configuration of the e-Genius-Mod is now available for more extensive flight tests and offers ideal conditions for scaled flight tests of all kinds.

Author Contributions: Conceptualization, D.P.B.; methodology, D.P.B., J.D., and O.P.; investigation, D.P.B., J.D., and O.P.; writing—original draft preparation, D.P.B., and J.D.; writing—review and editing, D.P.B., J.D., O.P., S.N., and A.S.; supervision, W.F., and A.S.; project administration, D.P.B.; funding acquisition, D.P.B., S.N., W.F., and A.S. All authors have read and agreed to the published version of the manuscript. Funding: This research was supported by Federal Ministry for Economic Affairs and Energy on the basis of a decision by the German Bundestag. Project: ELFLEAN—Electric wing tip propulsion system for the development of energy-efficient and noise reduced airplanes (20E1706). Institutional Review Board Statement: Not applicable. Informed Consent Statement: Not applicable. Data Availability Statement: Data not yet available public. Conflicts of Interest: The authors declare no conflict of interest.

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