AIAA 2012 147 188 Solstice


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Student paper on the hybrid propulsion system for UAV

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AIAA 2012 147 188 Solstice

  1. 1. 50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition AIAA 2012-014709 - 12 January 2012, Nashville, Tennessee SOLSTICE: Standalone-electric Optimized Lifting System, Transitional Internal Combustion Engine M. Cui, T. Drake, A. Kreuter, G. Kutil, B. Miller, C. Packard, M. Rahimpour and G. Soin 1 Aerospace Engineering Sciences, University of Colorado, Boulder, Colorado, 80309 This paper will focus on the innovative design architecture of the SOLSTICE hybrid aircraft propulsion system. The goal of this project is to design, manufacture, integrate and test a hybrid combustion/electric propulsion system to be integrated with a blended wing- body aircraft currently being designed by Hyperion, an international team led by graduate students at the University of Colorado. The concepts of operation will be discussed as well in order to elaborate on the development of a transitionable power control system, allowing the coaxial combustion engine and electric motor to operate in tandem or separately. Based on requirements provided by the Hyperion graduate team, the propulsion system is designed in part around power and thermal requirements. Analytical modeling of the engine system with respect to these facets has been performed and multiple testing scenarios will be run in order to validate these models. Verification of the hybrid propulsion system will be performed on the ground, setting flight testing on the Hyperion aircraft as a stretch goal. The SOLSTICE team plans to provide its customer, Dr. Jean Koster, with design recommendations for future hybrid aircraft propulsion endeavors, moving aircraft propulsion into a greener, fuel economy-friendlier realm. Nomenclature DAQ = data acquisition system EM = electric motor ESC = electronic speed controller HPS = hybrid propulsion system ICE = internal combustion engine RPM = revolutions per minute I. Introduction T HE Standalone-electric Optimized Lifting System, Transitional Internal Combustion Engine, or SOLSTICE, represents a turning point in how aircraft will be flown in the future. The goal of this project is to design, manufacture, integrate and test a hybrid combustion/electric propulsion system specifically for small aircraft. Currently, hybrid propulsion systems for aircraft are nearly nonexistent. However, the possibilities for their uses in multiple fields of aviation would allow for many performance benefits. An aircraft that is able to utilize a standard combustion system during normal flight and then transition to a much quieter electric motor during landing would greatly alleviate noise pollution, rampant at and around today’s airports. A transitional hybrid engine would also allow for greater fuel Figure 1. Hyperion Blended Wing-Body Aircraft Design. 1 Undergraduate Students, Aerospace Engineering Sciences Dept., Engineering Center 1111 Engineering Dr, AIAA Student Members 1 American Institute of Aeronautics and AstronauticsCopyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
  2. 2. efficiency (similar to a hybrid car) by decreasing fuel expenditure, allowing the combustion engine to operate moreefficiently while the electric motor provides any remaining power necessary for flight. The SOLSTICE HPS is one facet of a larger program known as Hyperion. The Hyperion team is an internationalamalgamation, comprised of students from the University of Stuttgart in Germany, from the University of Sydney inAustralia, and from the University of Colorado, represented by both undergraduate and graduate teams. Under theleadership of the CU graduate team, Hyperion is designing a model blended wing-body aircraft, seen in Fig. 1, withthe intent of developing a cleaner, greener and quieter aircraft system. The SOLSTICE undergraduate team isresponsible for the propulsion system of this new aircraft concept. The Hyperion team plans to fly the modeled aircraft in multiple flight modes, including cruise, quiet and dash,through the utilization of the SOLSTICE hybrid engine design. To achieve these different operational modes, theSOLSTICE engine controls two motors within the hybrid system either individually or in tandem: dash modeutilizes both the internal combustion engine and the electric motor while quiet mode uses the electric motor only.Cruise uses the combustion engine alone. SOLSTICE is designing the control system to operate these flight modesalong with the physical HPS. II. System Configuration The SOLSTICE hybrid propulsion system makes use of two separate engines, one electric motor (EM) and oneinternal combustion engine (ICE) to drive the propeller. A patent-pending gearbox converts the input torques ofthese two engines into one output torque, turning the propeller shaft. Power sources for the EM and ICE consist ofbatteries and fuel, respectively. By utilizing the EM and ICE either concurrently or independently the various flightmodes can be achieved such as EM only for quiet operation or EM and ICE concurrently for maximum power. Theentire system, excluding powersupplies, is placed on a singlebase plate to allow for ease ofintegration and structuralrigidity. The overall system isconstrained in mass and volumeas it is necessary for the HPS tofit inside the Hyperion aircraft.During the fall semester designphase, multiple analytical modelswere produced (thermal,structural stress and strain, andpower) in order to ensure that thesystem will provide thenecessary power to successfullyfly the Hyperion aircraft. Figure2 illustrates the propulsion Figure 2. SOLSTICE Hybrid Propulsion System Configuration.system configuration to bemanufactured in the spring semester, 2011. This configuration was designed iteratively over the course of the fall semester utilizing a systems engineeringapproach, working from broad to specific requirements. Initially the team worked with advisors and clients todevelop a preliminary goal and set of objectives. As is common practice, top level requirements were devised basedon these objectives and on certain parameters such as the overall aircraft mass restrictions. During the PreliminaryDesign phase, these requirements were broken down to a subsystem level where enough detail was defined in orderto narrow down selection of components. Trade studies were performed and the design continued with thoseselected. The overall system architecture was continually observed, allowing for a constant systems engineeringperspective. The results of these studies and their impact on the design were presented to a panel of aerospaceengineering faculty in a Preliminary Design Review (PDR). In the following Critical Design phase, parts wereselected and prototyping of major risks in order to find mitigating solutions was performed. During the CriticalDesign Review (CDR) modifications to the preliminary system architecture were presented by the team as a part ofthe complete system, its components and the analysis behind their design/selection. 2 American Institute of Aeronautics and Astronautics
  3. 3. This educational process goes hand-in-hand with aerospace industry practices. Project definition is the first andforemost step to the design process even with the existence of a multitude of unknowns at the time. This process-and the process to follow in the spring semester- has invaluably prepared the SOLSTICE team for any futureengineering endeavors. III. System Modeling Modeling physical and conceptual systems is a necessary process within engineering. The ability to predictperformance increases the probability of successfully completing a project. Prediction through modeling canproduce a more thorough understanding of a system’s characteristics and interdependencies. In most situations,modeling, combined with testing, generates this understanding earlier and more efficiently than testing alone.Increasing efficiency not only allows a project to be completed in a shorter period of time, but also decreases theamount of funds necessary for project completion.A. Power Modeling SOLSTICE’s main requirement from Hyperion, derived from the desired multiple flight modes, is that eitherthe electric motor or the internal combustion engine must supply the aircraft with 2 horsepower. During the design phase of the project, a simple, yet effective, power model was developed to ensure SOLSTICE’s HPS could meet Hyperion’s requirement. The model takes into account all major functional inefficiencies to predict the propulsion system’s power Figure 3. Power Model Component Efficiency. output. The major functional inefficiencies within the systeminclude friction losses from gearing and losses within the electric motor linked to converting electrical power tomechanical power. Figure 3 represents the structure of the model visually. It is predicted through analysis that thegearing is 75% efficient with a confidence of70%. The confidence is based on the possiblerange of coefficients of friction for meshinggears. The electric motor was assigned anefficiency of 80% based on manufacturerspecifications and lessons learned fromprevious experience. One of the novelties of the SOLSTICE Figure 4. Aircraft Flight Modes in Concept of Operations.engine is the capability of running ondifferent modes. This ultimately allows the aircraft to also fly at different flight modes and can be used to optimizeand limit fuel consumption. An example of how the variable flight operations can be utilized in an aircraft isshown in Fig. 4. Figure 4 shows how an aircraft can take advantage of the engine to achieve periods of maximum velocity,minimal fuel consumption, and regular flight operations. This concept was the foundation of the project: toconceive and build an engine capable of performing each of these flight modes. Each flight mode has a power output necessary to fly the Table 1. Required Propeller Output for Flight Modes aircraft as desired. The Flight Mode Power (W) Velocity (m/s) Ideal Prop RPM power output needed for Takeoff 1500 >7.6 1795.3 each flight mode is Cruise 750 24.3 5740.2 summarized in Table 1. Quiet 410 13.6 3212.6 The output power is the Dash 1500 32.4 7653.5 power required from the Landing 460 9.9 2338.6 propeller. 3 American Institute of Aeronautics and Astronautics
  4. 4. B. Thermal Modeling Of major concern in integration of the hybrid propulsion system with the aircraft is the generation of heat by the system. Because the propulsion system is located within the fuselage of the aircraft, the heat generated by the system will heat the surrounding aircraft structure and skin. The skin of the aircraft is fabricated from a fiberglass-epoxy material that must remain below 60oC in order to avoid material softening. To determine the thermodynamic characteristics of the system, two assumptions were made. The first assumption is that all power lost from inefficiencies in the system is fully converted to heat. The second assumption is no force convection over the system, only free convection. These two assumptions provide a worst case scenario of the propulsion system during operation. Through lumped analysis, it was determined that the system would produce 700W of heat via the three mechanisms of heat transfer. The heat transfer present within the engine cavity is shown in Fig. 5. In order to mitigate the risk of weakening the fiberglass epoxy’s integrity, forced convection over the system will be required. This shall be accomplished by implementing a duct into the nose of the aircraft. Further analysis of the system showed that a relation between the necessary forced convection properties, such as the convection Figure 5. Heat Transfer Model. coefficient and the flow speed, could be determined as a function of the total heat transfer. The relation is shown in Fig. 6 which gives the convection coefficient and flow speed required to keep fiberglass skinunder 60oC. The model proved that as the heat transfer decreases, the required airflow over the system decreases. It wasdetermined that a heat transfer of 700W is the worst case scenario thus the design team is confident that passive aircooling from the Hyperion aircraft will be sufficient for operational temperature. 150 Forced Convection Model 17 Convection Coefficient, h [W/m 2 K] Airflow Speed, V [m/s] 16 15 100 Mass Flow Rate (g/s) 14 13 50 12 11 10 0 0 100 200 300 400 500 600 700 9 270 275 280 285 290 295 300 Heat Transfer Q [W] Anticipated Ambient Air Temperature (K) Figure 6. Forced Convection Required. Figure 7. Necessary System Cooling. Further modeling has provided a mass flow rate necessary for operation, based on expected atmosphericconditions and the required interior temperature limit of 60 oC. The model predicts a mass flow rate range of 9.5 to16.5 g/s. Currently a mass flow rate equal to or greater than 34 g/s is expected to account for the worse casescenario and a factor of safety of 2. Figure 7 displays the mass flow rate’s dependence on the anticipatedfreestream temperature range. The model assumes steady state flow, the air is perfectly mixed within the system’scavity, and heat transfer is limited to convection within the cavity and conduction through the aircraft’s skin. 4 American Institute of Aeronautics and Astronautics
  5. 5. IV. System Verification The next step following system modeling is performing tests with real hardware and software. The dataobtained from system and subsystem tests can then be analyzed and used to verify the different system models. Inorder for these tests to represent the models accurately, the hardware including electronics and software usedduring the tests should be the actual flight hardware and not prototypes. System verification tests are the final stepbefore the overall system design can be validated and approved for flight. As such, it is essential that these tests beperformed following detailed laid out procedures as well as be repeated multiple times to obtain accurate data.A. Power Testing One of the primary requirements that the SOLSTICE HPS has to supply a minimum of 2 horsepower from its Electric Motor and Internal Combustion engine. These power requirements are essential for the Hyperion aircraft to achieve flight and also to verify their Concept of Operations. The SOLSTICE team performed tests on its HPS and used the data to verify the power output from the system. The primary apparatus used for these tests was a reaction force dynamometer which utilizes a force transducer to measure the reaction torque from the HPS. This is coupled with a voltage expander and an RPM sensor which together measure the input power to the system. The data obtained from these two sensors are then analyzed and compared in order to obtain the efficiency curves for the different components of the HPS. The setup for this apparatus can be seen in Fig. 8. 100 95 SOLSTICE EM Efficiency [%] Figure 8. Dynamometer with EM 90 Attached. Figures 9, 10 and 11 represent the efficiency 85and power curves for the most criticalcomponents of SOLSTICEs HPS namely the 80Electric Motor and the propeller. Using thedynamometer, SOLSTICE has been able to 75quantify the efficiency of both these EM Efficiencycomponents. Polyfit 70 Figure 9 shows the efficiency curve for the 4000 4500 5000 5500 6000 6500 7000 7500SOLSTICE Electric Motor plotted against its RPMRPM. The test was performed between a range Figure 9. EM Efficiency vs. RPM.of 5 to 20 Amps at intervals of 5 Amps in orderto replicate the desired flight conditions for the EM. The design specifications for the EM provided by themanufacturer rate the efficiency to be approximately 80%. From the EM dynamometer test, this specification wasverified and the efficiency of the motor was approximated to be 85% under the desired flight operation conditions. In order to verify that the SOLSTICE Electric Motor has the required power to satisfy the given powerrequirement, another dynamometer test was performed to quantify the power output of the motor. This was againplotted against RPM and is shown in Fig. 10. Due to structural limitations of the SOLSTICE Dynamometer, the testcould not be performed at the optimum range of input current for the EM. However, the test performed provided therelation through which the SOLSTICE team could then extrapolate the power output at higher current values.Through these analysis, it was determined that at an RPM of 8750, the SOLSTICE EM will output 1500 Watts ofpower which includes the 85% EM efficiency determined from the previous test. This power output is equivalent tothe 2 horsepower requirement. Hence, with the help of this test and its analysis, SOLSTICE has been able tosuccessfully verify its power requirements. 5 American Institute of Aeronautics and Astronautics
  6. 6. 800 Although, the power requirements give to SOLSTICE by the Hyperion 700 team accounted for a propeller efficiency of 50%, SOLSTICE did SOLSTICE EM Output Power 600 another dynamometer test to verify the efficiency of the actual flight propeller. 500 This was done in order to be thorough and obtain accurate power curves for 400 the HPS. Figure 11 shows the efficiency 300 curve of the flight propeller plotted against RPM along with a least squares 200 regression curve denoted as "Polyfit". The profile changes because 100 SOLSTICE is using a single speed 4000 4500 5000 5500 6000 6500 7000 7500 propeller which is designed to have a RPM Figure 10. EM Output Power in Watts(Including EM Efficiency) maximum efficiency at a single point. vs. RPM.The approximate propeller efficiency 70determined from this test is 55% to 65%which is higher than what the Hyperion 60requirements accounted for. Thus, the analysisfor this test further strengthens the verification 50of the SOLSTICE power requirements. Efficiency [%] 40 30 20 10 Propeller Efficiency Polyfit 0 1500 2000 2500 3000 3500 4000 4500 5000 5500 6000 RPM Figure 11. Propeller Efficiency vs. RPM.B. Thermal Testing To verify the thermodynamics and heat transfer of the hybrid propulsion system, tests are to be conducted toverify the requirements set forth by the structural integrity of the aircraft. The tests include measuring the thermal output from each component in the system. The major components that must be considered from a thermodynamic perspective are the internal combustion engine, electric motor, electronic speed controller, gearbox, and electronic circuits. Each component has inefficiencies where power losses generate heat. To experimentally determine the heat transfer of the system components, a test box is utilized that has the same geometry as the engine cavity within the aircraft. The test box also utilizes a polycarbonate material that has similar thermal properties as the fiberglass skin of the aircraft. A representation of the test box is shown in Fig. 12. In addition, the location of the temperature Figure 12. Thermal Test Bed Setup. sensors placed throughout the box to measure the 6 American Institute of Aeronautics and Astronautics
  7. 7. change in ambient temperature is shown. For the tests, seven LM 34 temperature sensors are to be used thatmeasure the ambient temperature in Fahrenheit. Using the Fahrenheit sensors allows for a better signal to noiseratio than that of the Celsius LM35 sensors. The first thermal test performed gave insight into how hot the engine cavity would get during the taxi portion ofthe flight. To simulate this, the internal combustion engine is to run on idle with the engine completely sealedwithin the box. During taxi, the propeller wash is negligible and is considered zero to test a worst case scenario.The test results, shown in Fig. 13, proved that over a span of 17 minutes, the temperature throughout the enginecavity does increase with the engine in idle. The maximum temperature measured during the 17 minute idle was40oC located directly above the idling engine.This is still below the maximum 60oC required be 45 Sensor 1the aircraft structure. When the aircraft takes to Sensor 2 40the air, the propeller will generate prop wash that Sensor 3will flow over the system and provide passive air Sensor 4 35 Sensor 5cooling. Ambient Temperature [C] Sensor 6 Testing of the electric motor and electronic Sensor 7 30speed controller are still to be performed howeverfrom preliminary dynamometer testing, the team 25is confident that these component will notgenerate any substantial heat and any forced 20convection will be sufficient for the requiredcooling. In addition, the gearbox will uselubrication that has the appropriate viscosity that 15will ensure minimal heat generation. Finally, afull system test is to be performed several times 10 0 2 4 6 8 10 12 14 16 18 20that will follow the aircraft mission and concept Time [min]of operations. This test will be performed a Figure 13. ICE Idle Results, No Forced Convection.multitude of times to ensure reliability in thethermodynamics and heat transfer of the hybrid propulsion system to the aircraft. V. System Applications The advantage of the SOLSTICE system comes from the ability to change between fundamentally differenttypes of engines. The ICE has the advantage of burning fuel, reducing the overall weight of the aircraft over time.The EM has the advantages of quiet operation relative to the ICE and it has the ability to run at higher altitudes. A typical use of the SOLSTICE engine will cycle the ICE and EM based on the flight regimes. The flightregimes include a dash mode which utilizes the power of both engines to reach a maximum speed for the aircraft.This engine mode also has the ability to provide significant thrust for takeoff through running the two enginestogether. This allows the selected engines to be smaller and tailored more towards the maximum thrust needed incruise conditions. The next flight regime utilizes the EM only. This flight regime can be considered a quiet modeand can be used in urban areas where noise levels are an issue. This mode can be used for unobtrusive surveillanceof targets as well. It has also been proposed that this flight regime be used as a high altitude mode, allowing theaircraft to fly at higher altitudes than those at which an ICE is capable of operating. The ICE-only flight regime canbe thought of as a general purpose mode. By burning the majority of the stored fuel during takeoff and climb thethrust needed in cruise would be minimized for a higher percent of the mission duration. In the future the ICE maybe made to use a bio-fuel such as biodiesel to decrease its environmental impact. This could potentially lower thefuel cost of operations which would be further offset by the use of the electric motor. The basic design of the SOLSTICE engine is not limited only to a propulsion system. A further applicationwould be as a portable generator for remote operations such as Antarctica or disaster relief. This would involvereplacing the EM with a generator and the propeller with a turbine blade optimized for power generation. Thesystem provides near-limitless power from the wind. This is augmented with the ICE as a traditional gas generator.By relying on the gas generator only when there is no wind the operation would need to carry less fuel for theirpower needs. In windy places the ICE could be seen as an emergency option. The combination of the two sourceswould reduce the total weight of the system and make it more portable than the two systems individually. 7 American Institute of Aeronautics and Astronautics
  8. 8. VI. Conclusion Although the application hybrid technology in aircraft has only recently become a topic of serious interest in theaerospace community, it is a necessary technology in ensuring that the aircraft of the future are both fuel efficientand environmentally friendly. The SOLSTICE project provides further development in the field of hybrid aircraftpropulsion through testing and modeling to verify the validity of such a system in aircraft. Current applications ofthe SOLSTICE project are limited to UAV markets due to current battery technologies. However, with an everincreasing interest in both UAV application and lessening carbon foot-prints, the SOLSTICE system is relevant intoday’s aircraft industry. The current state of testing verifies that the EM meets the power requirements and that thepropeller exceeds expected efficiency. Through the design process of the HPS, many valuable lessons have beenlearned including the importance of communication within the team and with the customer. Since the projectrepresents a global effort between students from various countries it is necessary that communication is clear andconcise to ensure that any problems that may arise are mitigated diligently and efficiently. By developing a systemwith the potential to rely less on fossil fuels for energy this project represents a global interest in ensuring that thefuture of aircraft propulsion is cleaner and safer. Acknowledgments The authors thank Dr. Jean Koster and Dr. Donna Gerren for their continued advice on this project. The teamalso appreciates the strong support and guidance received from Cody Humbargar and the entire Hyperion team(Scott Balaban, Derek Nasso, Andrew Brewer, Julie Price, Chelsea Goodman, Eric Serani, Derek Hillery, AlecVelazco, Mark Johnson, Richard Zhao, and Tom Wiley). We also thank the international partner teams from theUniversity of Stuttgart in Germany and from the University of Sydney in Australia who helped make this project afascinating and unique learning experience. Funding for this project is provided in part by The Boeing Company,eSpace Inc., and NASA under grant NNX09AF65G. 8 American Institute of Aeronautics and Astronautics