James L. Felder

Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

Meeting NASAs N+3 goals requires a fundamental shift in approach to aircraft and engine design. Material and design improvements allow higher pressure and higher temperature core engines which improve the thermal efficiency. Propulsive efficiency, the other half of the overall efficiency equation, however, is largely determined by the fan pressure ratio (FPR). Lower FPR increases propulsive efficiency, but also dramatically reduces fan shaft speed through the combination of larger diameter fans and reduced fan tip speed limits. The result is that below an FPR of 1.5 the maximum fan shaft speed makes direct drive turbines problematic. However, it is the low pressure ratio fans that allow the improvement in propulsive efficiency which, along with improvements in thermal efficiency in the core, contributes strongly to meeting the N+3 goals for fuel burn reduction. The lower fan exhaust velocities resulting from lower FPRs are also key to meeting the aircraft noise goals. Adding a gear box to the standard turbofan engine allows acceptable turbine speeds to be maintained. However, development of a 50,000+ hp gearbox required by fans in a large twin engine transport aircraft presents an extreme technical challenge, therefore another approach is needed. This paper presents a propulsion system which transmits power from the turbine to the fan electrically rather than mechanically. Recent and anticipated advances in high temperature superconducting generators, motors, and power lines offer the possibility that such devices can be used to transmit turbine power in aircraft without an excessive weight penalty. Moving to such a power transmission system does more than provide better matching between fan and turbine shaft speeds. The relative ease with which electrical power can be distributed throughout the aircraft opens up numerous other possibilities for new aircraft and propulsion configurations and modes of operation. This .

An Examination of the Effect of Boundary Layer Ingestion on Turboelectric Distributed Propulsion Systems

A Turboelectric Distributed Propulsion (TeDP) system differs from other propulsion systems by the use of electrical power to transmit power from the turbine to the fan. Electrical power can be efficiently transmitted over longer distances and with complex topologies. Also the use of power inverters allows the generator and motors speeds to be independent of one another. This decoupling allows the aircraft designer to place the core engines and the fans in locations most advantageous for each. The result can be very different installation environments for the different devices. Thus the installation effects on this system can be quite different than conventional turbofans where the fan and core both see the same installed environments. This paper examines a propulsion system consisting of two superconducting generators, each driven by a turboshaft engine located so that their inlets ingest freestream air, superconducting electrical transmission lines, and an array of superconducting motor driven fan positioned across the upper/rear fuselage area of a hybrid wing body aircraft in a continuous nacelle that ingests all of the upper fuselage boundary layer. The effect of ingesting the boundary layer on the design of the system with a range of design pressure ratios is examined. Also the impact of ingesting the boundary layer on offdesign performance is examined. The results show that when examining different design fan pressure ratios it is important to recalculate of the boundary layer mass-average Pt and MN up the height for each inlet height during convergence of the design point for each fan design pressure ratio examined. Correct estimation of off-design performance is dependent on the height of the column of air measured from the aircraft surface immediately prior to any external diffusion that will flow through the fan propulsors. The mass-averaged Pt and MN calculated for this column of air determine the Pt and MN seen by the propulsor inlet. Since the height of this column will change as the amount of air passing through the fans change as the propulsion system is throttled, and since the mass-average Pt and MN varies by height, this “capture height” must be recalculated as the airflow through the propulsor is varied as the off-design performance point is converged.

Sensitivity of Mission Energy Consumption to Turboelectric Distributed Propulsion Design Assumptions on the N3-X Hybrid Wing Body Aircraft

In a previous study by the authors it was shown that the N3-X, a 300 passenger hybrid wing body (HWB) aircraft with a turboelectric distributed propulsion (TeDP) system, was able to meet the NASA Subsonic Fixed Wing (SFW) project goal for N+3 generation aircraft of at least a 60% reduction in total energy consumption as compared to the best in class current generation aircraft. This previous study combined technology assumptions that represented the highest anticipated values that could be matured to technology readiness level (TRL) 4-6 by 2030. This paper presents the results of a sensitivity analysis of the total mission energy consumption to reductions in each key technology assumption. Of the parameters examined, the mission total energy consumption was most sensitive to changes to total pressure loss in the propulsor inlet. The baseline inlet internal pressure loss is assumed to be an optimistic 0.5%. An inlet pressure loss of 3% increases the total energy consumption 9%. However changes to reduce inlet pressure loss can result in additional distortion to the fan which can reduce fan efficiency or vice versa. It is very important that the inlet and fan be analyzed and optimized as a single unit. The turboshaft hot section is assumed to be made of ceramic matrix composite (CMC) with a 3000 F maximum material temperature. Reducing the maximum material temperature to 2700 F increases the mission energy consumption by only 1.5%. Thus achieving a 3000 F temperature in CMCs is important but not central to achieving the energy consumption objective of the N3-X/TeDP. A key parameter in the efficiency of superconducting motors and generators is the size of the superconducting filaments in the stator. The size of the superconducting filaments in the baseline model is assumed to be 10 microns. A 40 micron filament, which represents current technology, results in a 200% increase in AC losses in the motor and generator stators. This analysis shows that for a system with 40 micron filaments the higher stator losses plus the added weight and power of larger cryocoolers results in a 4% increase in mission energy consumption. If liquid hydrogen is used to cool the superconductors the 40 micron fibers results in a 200% increase in hydrogen required for cooling. Each pound of hydrogen used as fuel displaces 3 pounds of jet fuel. For the N3-X on the reference mission the additional hydrogen due to the increase stator losses reduces the total fuel weight 10%. The lighter fuel load and attendant vehicle resizing reduces the total energy consumption more than the higher stator losses increase it. As a result with hydrogen cooling there is a slight reduction in mission energy consumption with increasing stator losses. This counter intuitive result highlights the need to consider the full system impact of changes rather than just at the component or subsystem level.

Turboelectric Aircraft Drive Key Performance Parameters and Functional Requirements

The purpose of this paper is to propose specific power and efficiency as the key performance parameters for a turboelectric aircraft power system and investigate their impact on the overall aircraft. Key functional requirements are identified that impact the power system design. Breguet range equations for a base aircraft and a turboelectric aircraft are found. The benefits and costs that may result from the turboelectric system are enumerated. A break-even analysis is conducted to find the minimum allowable electric drive specific power and efficiency that can preserve the range, initial weight, operating empty weight, and payload weight of the base aircraft.

Control Volume Analysis of Boundary Layer Ingesting Propulsion Systems With or Without Shock Wave Ahead of the Inlet

The performance benefit of boundary layer or wake ingestion on marine and air vehicles has been well documented and explored. In this article, a quasi-one-dimensional boundary layer ingestion (BLI) benefit analysis for subsonic and transonic propulsion systems is performed using a control volume of a ducted propulsion system that ingests the boundary layer developed by the external airframe surface. To illustrate the BLI benefit, a relationship between the amount of BLI and the net thrust is established and analyzed for two propulsor types. One propulsor is an electric fan, and the other is a pure turbojet. These engines can be modeled as a turbofan with an infinite bypass ratio for the electric fan, and with a zero bypass ratio for the pure turbojet. The analysis considers two flow processes: a boundary layer being ingested by an aircraft inlet and a shock wave sitting in front of the inlet. Though the two processes are completely unrelated, both represent a loss of total pressure and velocity. In real applications, it is possible to have both processes occurring in front of the inlet of a transonic vehicle. Preliminary analysis indicates that the electrically driven propulsion system benefits most from the boundary layer ingestion and the presence of transonic shock waves, whereas the benefit for the turbojet engine is near zero or negative depending on the amount of total temperature rise across the engine.

Overview of NASA Electrified Aircraft Propulsion Research for Large Subsonic Transports

NASA is investing in Electrified Aircraft Propulsion (EAP) research as part of the portfolio to improve the fuel efficiency, emissions, and noise levels in commercial transport aircraft. Turboelectric, partially turboelectric, and hybrid electric propulsion systems are the primary EAP configurations being evaluated for regional jet and larger aircraft. The goal is to show that one or more viable EAP concepts exist for narrow-body aircraft and mature tall-pole technologies related to those concepts. A summary of the aircraft system studies, technology development, and facility development is provided. The leading concept for midterm (2035) introduction of EAP for a single-aisle aircraft is a tube and wing, partially turboelectric configuration NASA Single-Aisle Turboelectric Aircraft With Aft Boundary Layer (STARC– ABL); however, other viable configurations exist. Investments are being made to raise the technology readiness level of lightweight, high-efficiency motors, generators, and electrical power distribution systems as well as to define the optimal turbine and boundary-layer ingestion systems for a midterm tube and wing configuration. An electric aircraft power system test facility (NASA Electric Aircraft Testbed (NEAT)) is under construction at NASA Glenn Research Center and an electric aircraft control system test facility (Hybrid Electric Integrated System Testbed (HEIST)) is under construction at NASA Armstrong Flight Research Center. The correct building blocks are in place to have a viable large-plane EAP configuration tested by 2025 leading to entry into service in 2035 if the community chooses to pursue that goal.

Turboelectric distributed propulsion benefits on the N3-X vehicle

Purpose – The purpose of this article is to present a summary of recent study results on a turboelectric distributed propulsion vehicle concept named N3-X. Design/methodology/approach – The turboelectric distributed propulsion system uses multiple electric motor-driven propulsors that are distributed on an aircraft. The power to drive these electric propulsors is generated by separately located gas turbine-driven electric generators on the airframe. To estimate the benefits associated with this new propulsion concept, a system analysis was performed on a hybrid-wing-body transport configuration to determine fuel burn (or energy usage), community noise and emissions reductions. Findings – N3-X would be able to reduce energy consumption by 70-72 per cent compared to a reference vehicle, a Boeing 777-200LR, flying the same mission. Predictions for landing and take-off NOX are estimated to be 85 per cent less than the Tier 6-CAEP/6 standard. Two variants of the N3-X vehicle were examined for certification noi...

System Weight Comparison of Electric Machine Topologies for Electric Aircraft Propulsion

An important thrust in current aerospace research is aircraft electrification, including propulsion system electrification. For an electrified propulsion system to provide net benefit over conventional propulsion, high specific power, power density, and efficiency requirements of the electrical system must be met. This paper expands on previous work by comparing electric machine topologies for electric aircraft propulsors while considering tradeoffs in the power electronics, fault response equipment, and gearbox components. Permanent magnet synchronous machine (PMSM), brushless DC machine (BLDC), switched reluctance machine (SRM), brushless doubly-fed reluctance machine (BDFRM), and induction machine (IM) topologies are explored. A parametric design tool including finite element analysis is used to create a viable design for each machine type. Next, analytical sizing equations are used to scale the designed machines to varying operating speeds and aspect ratios. In addition, gearbox, power electronics, circuit breaker, clutch, thermal management system, and energy storage weights are predicted based on current power densities of research designs. It is well established that PMSMs have outstanding power density and this study shows that when considering additional drive and fault response components, PMSMs still maintain the weight advantage.