Fuel Burn Benefits Research Articles

Blended-Wing-Body Regional Aircraft Optimization with High-Fidelity Aerodynamics and Critical Design Requirements

Conventional tube-and-wing and proposed blended-wing-body airliners must satisfy several design requirements, but the latter configuration is tightly integrated and sensitive to these requirements. In this work, blended-wing-body regional aircraft are investigated using a gradient-based mixed-fidelity multidisciplinary optimization framework centered on a Reynolds-averaged Navier–Stokes solver. In addition to sizing and cruise trim, design requirements considered include one-engine-inoperative directional trim, takeoff rotation ability, takeoff field length, initial climb performance, low-speed trim and static margin, and top-of-climb rate of climb. Results show how optimal design features vary and performance is overpredicted if critical design requirements are excluded and how key elements of geometric freedom help realize the potential of the configuration. A 4.8% block fuel burn benefit is enabled by pivot-piston variable-length landing gear even with low-mounted engines and high design freedom. The off-design constraints penalize block fuel burn by 3.2% if variable-length landing gear is considered, but this value reaches 7.6% if lower geometric freedom that inhibits tight cabin contouring and the formation of a novel forebody ridge is removed. Leading-edge carving is found to be optimal. The high design freedom and high-fidelity aerodynamics model help efficiently satisfy the design requirements, resulting in a cruise lift-to-drag ratio of 21.7 at 36,000 ft and Mach 0.78.

Economic Aspects of Aircraft Propulsion Electrification

Aircraft propulsion electrification is currently being considered by industry and academia as one of the most promising strategies to reduce air transport emissions and increase overall efficiency levels. In the past decade, several papers were published on this subject, with the majority indicating encouraging fuel burn benefits versus conventional, fossil-fuel-based propulsion systems when future technologies, novel aircraft configurations, and synergistic propulsive-airframe integration are employed. However, a much smaller effort has been applied to the economic aspects of hybrid and fully electric propulsion, which are crucial for a successful product introduction. The present paper describes the modeling of a baseline general-aviation-type aircraft and its propulsion system retrofit with electrified architectures, exploring different electrification strategies for a fixed airframe design. Analyses are performed at the aircraft level, comparing recurring and cash operating costs for several cost and durability scenarios. While considerable CO2 reductions may be achieved in some electrification strategies, aircraft performance is significantly penalized, and important improvements in economic figures of merit are needed in order to make electrified propulsion cost-competitive. Electrified architectures tend to increase costs: turboelectric increases recurring equipment costs, while hybrid-electric increases recurring and direct maintenance costs, especially at higher degrees of energy hybridization.

Integrated Hybrid Engine Cycle Design and Power Management Optimization

Abstract A novel integrated gas turbine cycle design and power management optimization methodology for parallel hybrid electric propulsion architectures is presented in this paper. The gas turbine multipoint cycle design method is extended to turboprop and turbofan architectures, and several trade studies are performed initially at the cycle level. It is shown that the maximum degree of electrification is limited by the surge margin levels of the booster in the turbofan configuration. An aircraft mission-level assessment is then performed using the integrated optimization method initially for an A320 Neo style aircraft case. The results indicate that the optimal cycle redesigned hybrid electric propulsion system (HEPS) favors takeoff and climb power on-takes while optimal retrofit HEPS favor cruise power on-takes. It is shown that for current battery energy density (250 Wh/Kg), there is no fuel burn benefit. Furthermore, even for optimistic energy density values (750 Wh/kg) the maximum fuel burn benefit for a 500 nm mission is 5.5% and 4% for redesigned and retrofit HEPS, respectively. The power management strategies for HEPS configurations also differ based on gas turbine technology with improvement in gas turbine technology showing greater scope for electrification. The method is then extended to ATR 72 style aircraft case, showing greater fuel burn benefits across the flight mission envelope. The power management strategies also change depending on the objective function, and optimum strategies are reported for direct operating cost or fuel burn. The retrofit case studies show a benefit in direct operating cost compared to redesigned case studies for ATR 72. Finally, a novel multimission approach is shown to highlight the overall fuel burn and direct operating cost benefit across the aircraft mission cluster.

Blended Wing Body Configuration for Hydrogen-Powered Aviation

Hydrogen aircraft have the potential to achieve more significant climate impact reductions at a lower cost than aircraft powered by biofuels or other drop-in sustainable aviation fuels. Nevertheless, even as a liquid, hydrogen requires four times more volume than kerosene to store the same energy. Companies and researchers have suggested that the blended wing body configuration is well-suited to hydrogen because it can efficiently store large fuel tanks. However, nobody has quantified this claim, at least publicly. We address this gap by comparing optimized kerosene and hydrogen versions of blended wing body and tube and wing aircraft. Our models predict that, with ambitious hydrogen tank technology assumptions, a hydrogen blended wing body has a 3.8% energy penalty compared to a kerosene blended wing body. In contrast, the energy penalty for a tube and wing is 5.1%. These results suggest that the blended wing body configuration is slightly better suited to hydrogen than a tube and wing. However, the conclusion breaks down under less optimistic scenarios where lightweight conformal liquid hydrogen tanks are unavailable. Nonetheless, the energy penalties of switching the two configurations to hydrogen have a similar order of magnitude. Any kerosene blended wing body fuel burn benefit over a tube and wing also applies to hydrogen versions. As a result, kerosene blended wing body technology developments, such as maturing the structural design of the noncircular pressurized passenger cabin, remain valuable for future hydrogen-powered aircraft.

Off-Design Performance Analysis of a Parallel Hybrid Electric Regional Turboprop Aircraft

Hybrid electric propulsion is one of the alternative solutions for reducing fuel burn and lower emissions while keeping a reasonable battery mass for regional turboprop aircraft operating on short routes. Most studies reporting fuel burn reductions evaluate the aircraft on the design mission, although regional transport aircraft rarely operate under these conditions. Therefore, considering its off-design performance is essential for providing a more complete understanding of aircraft capabilities under various operating conditions. Under these flight conditions, the multi-energy management aspect of hybrid propulsion and the fixed size of the batteries could have a significant impact on the system robustness in off-design operation. In this study, the off-design performance of an existing regional turboprop aircraft retrofitted with a parallel hybrid electric powertrain is assessed. Fuel burn benefits are evaluated on the payload–range diagram for an initial hybrid design and compared to the baseline aircraft. Then, using a novel sizing approach, considering a typical mission operation, this study shows an average improvement of percentage point on fuel burn benefits relative to the initial hybrid aircraft, creating a more robust design.

Design Methodology and Mission Assessment of Parallel Hybrid Electric Propulsion Systems

Abstract An integrated engine cycle design methodology and mission assessment for parallel hybrid electric propulsion architectures are presented in this paper. The aircraft case study considered is inspired by Fokker 100, boosted by an electric motor on the low-pressure shaft of the gas turbine. The fuel burn benefits arising from boosting the low-pressure shaft are discussed for two different baseline engine technologies. A three-point engine cycle design method is developed to redesign the engine cycle according to the degree of hybridization. The integrated cycle design and power management optimization method is employed to identify potential fuel burn benefits from hybridization for multiple mission ranges. The sensitivity of these mission results has also been analyzed for different assumptions on the electric powertrain. With 1 MW motor power and a battery pack of 2307 kg, a 3% fuel burn benefit can be obtained by retrofitting the gas turbine for 400 nm range. Optimizing the power management strategy improves this fuel burn benefit by 0.2-0.3%. Redesigning the gas turbine and optimizing the power management strategy, provides a 4.2% fuel benefit on 400 nm. The results suggest that a high hybridization by power, low hybridization by energy, and ranges below 700 nm are the only cases where the redesigned hybrid electric aircraft has benefits in fuel burn and energy consumption relative to the baseline aircraft. Finally, it is found that the percentage of fuel burn benefits from the hybrid electric configuration increases with the improvement in engine technology.

System Analysis and Design Space Exploration of Regional Aircraft with Electrified Powertrains

This paper explores the design spaces of a thin-haul and a regional aircraft with parallel hybrid-electric propulsion architectures for a 2030 entry into service date. Notional technology reference aircraft models were developed for a 19- and a 50-passenger aircraft based on publicly available data on Beechcraft 1900D and ATR 42-600, respectively. A set of airframe and propulsion system technologies projected to reach maturity by 2030 was infused into the aircraft models. Parametric, physics-based models were created for the charge-depleting hybrid architecture. Different modes of operation were identified and parameterized with a range of design variables to investigate the feasibility and trade space for peak power shaving, climb power boosting, electric taxi, battery usage schedules, and in-flight battery recharge strategies. Thousands of electrified aircraft concepts with varying electrification, operation, and technology scenarios were sized under the same system-level requirements as their conventional counterpart. The resulting multidisciplinary design space exploration environment was used to identify the optimum vision system designs and modes of operation for the minimum block fuel-burn objective. It has been found that both vehicle classes with the charge-depleting parallel hybrid electric architecture provided fuel-burn benefits over their 2030 advanced technology counterparts under certain operational modes.

Fan Aerodynamics With a Short Intake at High Angle of Attack

Abstract Fans that are designed to maintain thrust at the high angle of attack (AOA) flight condition could exploit the cruise fuel burn benefit of a shorter intake design. This article considers how the fan rotor radial pressure ratio distribution and tip velocity triangle can be designed to improve thrust when coupled to a short intake operating at high AOA. Two AOA values are investigated using unsteady computational fluid dynamics: 20 deg (attached flow) and 35 deg (separated flow). Thrust at high AOA is governed by three key loss and work input mechanisms. (i) Rotor choking loss: flow is accelerated around the intake bottom lip and enters the rotor with high Mach numbers. Fans designed with a tip-high radial pressure ratio distribution reduced choking loss with a separated intake compared to a mid-high design, particularly when the tip velocity triangle was designed with high diffusion instead of high camber. (ii) Rotor–separation interaction loss: the rotor ingests low mass flow when operating inside the separation and the casing boundary layer separates. High diffusion tip designs strengthened the casing separation, but this penalty did not outweigh improved choking loss. (iii) Work input in radial flows: high AOA generates strong radial flows through the rotor, which alter both the amount and the way work is imparted on the flow. Fans designed with a mid-high radial pressure ratio distribution imparted high work on streamlines migrating toward the hub. Consolidating these findings, we propose two design philosophies for improved thrust at high AOA: high work (mid-high radial pressure ratio distribution) or low loss (tip-high radial pressure ratio distribution with high diffusion tip velocity triangle).

Novel Aero-Engine Multi-Disciplinary Preliminary Design Optimization Framework Accounting for Dynamic System Operation and Aircraft Mission Performance

This paper presents a modular, flexible, extendable and fast-computational framework that implements a multidisciplinary, varying fidelity, multi-system approach for the conceptual and preliminary design of novel aero-engines. In its current status, the framework includes modules for multi-point steady-state engine design, aerodynamic design, engine geometry and weight, aircraft mission analysis, Nitrogen Oxide (NOx) emissions, control system design and integrated controller-engine transient-performance analysis. All the modules have been developed in the same software environment, ensuring consistent and transparent modeling while facilitating code maintainability, extendibility and integration at modeling and simulation levels. Any simulation workflow can be defined by appropriately combining the relevant modules. Different types of analysis can be specified such as sensitivity, design of experiment and optimization. Any combination of engine parameters can be selected as design variables, and multi-disciplinary requirements and constraints at different operating points in the flight envelope can be specified. The framework implementation is exemplified through the optimization of an ultra-high bypass ratio geared turbofan engine with a variable area fan nozzle, for which specific aircraft requirements and technology limits apply. Although the optimum design resulted in double-digit fuel-burn benefits compared to current technology engines, it did not meet engine-response requirements, highlighting the need to include transient-performance assessments as early as possible in the preliminary engine design phase.

Modeling Geared Turbofan and Open Rotor Engine Performance for Year-2050 Long-Range and Short-Range Aircraft

Abstract This paper provides design and performance data for two envisaged year-2050 engines: a geared high bypass turbofan for intercontinental missions and a contra-rotating pusher open rotor targeting short to medium range aircraft. It defines component performance and cycle parameters, general arrangements, sizes, and weights. Reduced thrust requirements reflect expected improvements in engine and airframe technologies. Advanced simulation platforms have been developed to model the engines and details of individual components. The engines are optimized and compared with “baseline” year-2000 turbofans and an anticipated year-2025 open rotor to quantify the relative fuel-burn benefits. A preliminary scaling with year-2050 “reference” engines, highlights tradeoffs between reduced specific fuel consumption (SFC) and increased engine weight and diameter. These parameters are converted into mission fuel burn variations using linear and nonlinear trade factors (NLTF). The final turbofan has an optimized design-point bypass ratio (BPR) of 16.8, and a maximum overall pressure ratio (OPR) of 75.4, for a 31.5% TOC thrust reduction and a 46% mission fuel burn reduction per passenger kilometer compared to the respective “baseline” engine–aircraft combination. The open rotor SFC is 9.5% less than the year-2025 open rotor and 39% less than the year-2000 turbofan, while the TOC thrust increases by 8% versus the 2025 open rotor, due to assumed increase in passenger capacity. Combined with airframe improvements, the final open rotor-powered aircraft has a 59% fuel-burn reduction per passenger kilometer relative to its baseline.

Thermodynamic analysis of nutating disc engine topping cycles for aero-engine applications

Within the next thirty years the evolutionary approach to aero engine development will struggle to keep abreast with increasingly stringent environmental targets. Therefore radical approaches to aero-engine development in terms of energy savings need to be considered. One particular concept involves the inclusion of a pressure-rise combustion system, within the architecture of an aero-engine, to provide additional shaft power. The nutating disc engine concept is a strong contender due to its power density. The feasibility of the nutating disc engine has been previously investigated for unmanned vehicle applications. However, this paper investigates the performance benefits of incorporating a nutating disc core in a larger geared open rotor engine for a potential entry in to service in 2050. In addition, a methodology is presented to estimate the size and weight of the nutating disc core. This methodology is pivotal in determining the overall performance of the novel aero-engine cycle. The outcome of this study predicts a potential 9.4% fuel burn benefit, over a state of the art geared open rotor in the year 2050. In addition, the sensitivity of the nutating disc design variables highlights the possible fuel burn benefits compared against a comparable year-2000 aircraft mission.

Investigation on the benefit of formation flight with a focus on the leading and trailing aircraft rotation

PurposeThe purpose of this paper is to study the potential advantages of aircraft formation flight (FF) and to exploit further benefits through exchanging the leading positions.Design/methodology/approachThe detailed and robust methodologies concerning FF mission analysis including the leading aircraft rotation strategies are developed in this paper to study the fuel burn benefit and the additional bonus of formation rotation.FindingsSwitch of FF leading positions can offset the undesired weight ratios between the leading and trailing aircraft within FF missions, which further alleviates the deviations from design flight conditions. The case studies on two long-range civil transport aircraft in FF show that the leading and trailing aircraft can achieve almost equal fuel benefit through rotations. As compared to FF without rotation, the fuel efficiency can be improved by more than 11 per cent.Research limitations/implicationsThe work can bring benefit the research communities as a fundamental basis for operational studies of FF, such as FF airspace management in the future, which is significant for a future real-world implementation of FFs.Practical implicationsAccording to the authors’ study, equal or quasi-equal fuel savings can be achieved if the rotation is properly arranged. For the real-world FF application, fuel consumption (FC) or cost redistribution problem for leading and trailing aircraft belonging to two different operating airlines can therefore be resolved through the concept proposed by the paper.Originality/valueThe methods developed in the paper have the advantage to give more reliable estimations of the achievable fuel burn savings of FF. The concept proposed in the paper has significant meaning with respect to offset the undesired weight ratios between the leading and trailing aircraft within FF missions and redistributing FC or cost redistribution of different operating airlines.

Estimation of Vehicle-Level Fuel Burn Benefits of Aircraft Formation Flight

The present paper investigates the overall aircraft-level impact of formation flight in terms of total drag savings and fuel consumption reductions. The current work is carried out on the basis of an in-house multidisciplinary integrated conceptual aircraft design optimization framework. An aerodynamic method based on potential flow theory is first adapted for the induced drag estimation of aircraft in formation flight. By integrating the formation flight aerodynamic methodology into the in-house aircraft design and optimization platform, the potential drag reduction for arbitrary aircraft configurations in formation can be assessed. For specific flight missions, the fuel burn reduction can be further determined. Through comparing with other simulation and experimental results, the proposed formation flight aerodynamic methods in the present work show very good agreement. A case study on two medium-range passenger aircraft indicates that the trailing aircraft could achieve up to 60% induced drag reduction with optimal longitudinal, lateral, and vertical aircraft separations within formation. The fuel burn savings of the trailing aircraft are calculated to be up to 16% for specific payload and flight range combinations. A reserve fuel study shows that around 2% fuel burn savings would be eroded due to conservative fuel planning in formation flight.

Performance of a Supercritical CO2 Bottoming Cycle for Aero Applications

By 2050, the evolutionary approach to aero engine research may no longer provide meaningful returns on investment, whereas more radical approaches to improving thermal efficiency and reducing emissions might still prove cost effective. One such radical concept is the addition of a secondary power cycle that utilizes the otherwise largely wasted residual heat in the core engine’s exhaust gases. This could provide additional shaft power. Supercritical carbon dioxide closed-circuit power cycles are currently being investigated primarily for stationary power applications, but their high power density and efficiency, even for modest peak cycle temperatures, makes them credible bottoming cycle options for aero engine applications. Through individual geometric design and performance studies for each of the bottoming cycle’s major components, it was determined that a simple combined cycle aero engine could offer a 1.9% mission fuel burn benefit over a state-of-the-art geared turbofan for the year 2050. However, the even greater potential of more complex systems demands further investigation. For example, adding inter-turbine reheat (ITR) to the combined cycle is predicted to significantly improve the fuel burn benefit.

First and Second Law Analysis of Intercooled Turbofan Engine

Although the benefits of intercooling for aero-engine applications have been realized and discussed in many publications, quantitative details are still relatively limited. In order to strengthen the understanding of aero-engine intercooling, detailed performance data on optimized intercooled (IC) turbofan engines are provided. Analysis is conducted using an exergy breakdown, i.e., quantifying the losses into a common currency by applying a combined use of the first and second law of thermodynamics. Optimal IC geared turbofan engines for a long range mission are established with computational fluid dynamics (CFD) based two-pass cross flow tubular intercooler correlations. By means of a separate variable nozzle, the amount of intercooler coolant air can be optimized to different flight conditions. Exergy analysis is used to assess how irreversibility is varying over the flight mission, allowing for a more clear explanation and interpretation of the benefits. The optimal IC geared turbofan engine provides a 4.5% fuel burn benefit over a non-IC geared reference engine. The optimum is constrained by the last stage compressor blade height. To further explore the potential of intercooling the constraint limiting the axial compressor last stage blade height is relaxed by introducing an axial radial high pressure compressor (HPC). The axial–radial high pressure ratio (PR) configuration allows for an ultrahigh overall PR (OPR). With an optimal top-of-climb (TOC) OPR of 140, the configuration provides a 5.3% fuel burn benefit over the geared reference engine. The irreversibilities of the intercooler are broken down into its components to analyze the difference between the ultrahigh OPR axial–radial configuration and the purely axial configuration. An intercooler conceptual design method is used to predict pressure loss heat transfer and weight for the different OPRs. Exergy analysis combined with results from the intercooler and engine conceptual design are used to support the conclusion that the optimal PR split exponent stays relatively independent of the overall engine PR.

Parametric Analysis and Design for Embedded Engine Inlets

The present paper demonstrates the aerodynamic feasibility of boundary-layer ingesting embedded-engine inlet designs with low total pressure losses and distortion harmonic content. The inlet was designed using a hierarchical multi-objective computational fluid dynamics optimization that combined global and local shaping. Global parameters including duct offset and length, wall curvature and shape, inlet aspect ratio, lip contour and thickness, and upstream airframe contour were used to identify optimal design space regions. Local inlet shaping optimization further reduced total pressure losses and harmonic distortion upstream of the fan. Coupled inlet/fan design iterations were carried out for selected inlet designs to assess the fan/engine stability and operability benefits. The resulting inlet design has the potential of achieving a 3–5% boundary-layer ingesting fuel burn benefit for NASA’s Generation-After-Next aircraft relative to a baseline high-performance pylon-mounted propulsion system. It shows significantly improved performance when compared with NASA’s “inlet A” reference geometry with a length-to-diameter ratio of 3. The new inlet was shortened to a length-to-diameter ratio of 0.6, total pressure losses were reduced by three times, dominant distortion harmonic amplitudes were reduced by 30–50%, and fan efficiency losses were reduced from 6 to 0.5–1.5%. No flow control was required.

Surface Performance of End-Around Taxiways

Optimizing usage of end-around taxiways is a near-term operational change to reduce aviation emissions and increase efficiency at airports. An end-around taxiway is a path for an aircraft to taxi around an active runway. End-around taxiways provide benefits such as increased throughput and safety, reduced surface congestion, thus also yielding environmental benefits. This study analyzes end-around taxiway use at three airports: Atlanta (ATL), Dallas (DFW), and Detroit (DTW) using ASDE-X data over a six-month period. We developed three types of decision rules to maximize fuel savings. The most promising (environmentally beneficial) decision rule at each airport is based on multiple factors, including terminal destination and arrival time. Depending on the airport, the multi-factor rules resulted in an average aircraft taxi-in fuel savings of 8.9% to 25.4%. This research provides decision-makers at the operational level with a practical guidance tool to use end-around taxiways effectively. This research focuses on reducing taxi-in fuel burn; we do not directly consider the impact of the decision rules on departure parameters (such as throughput and taxi-out fuel burn). Future work will expand the model to optimize fuel burn benefits across integrated arrival and departure surface operations.

Ultrashort Nacelles for Low Fan Pressure Ratio Propulsors

As the propulsor fan pressure ratio (FPR) is decreased for improved fuel burn, reduced emissions and noise, the fan diameter grows and innovative nacelle concepts with short inlets are required to reduce their weight and drag. This paper addresses the uncharted inlet and nacelle design space for low-FPR propulsors where fan and nacelle are more closely coupled than in current turbofan engines. The paper presents an integrated fan–nacelle design framework, combining a spline-based inlet design tool with a fast and reliable body-force-based approach for the fan rotor and stator blade rows to capture the inlet–fan and fan–exhaust interactions and flow distortion at the fan face. The new capability enables parametric studies of characteristic inlet and nacelle design parameters with a short turn-around time. The interaction of the rotor with a region of high streamwise Mach number at the fan face is identified as the key mechanism limiting the design of short inlets. The local increase in Mach number is due to flow acceleration along the inlet internal surface coupled with a reduction in effective flow area. For a candidate short-inlet design with length over diameter ratio L/D = 0.19, the streamwise Mach number at the fan face near the shroud increases by up to 0.16 at cruise and by up to 0.36 at off-design conditions relative to a long-inlet propulsor with L/D = 0.5. As a consequence, the rotor locally operates close to choke resulting in fan efficiency penalties of up to 1.6% at cruise and 3.9% at off-design. For inlets with L/D < 0.25, the benefit from reduced nacelle drag is offset by the reduction in fan efficiency, resulting in propulsive efficiency penalties. Based on a parametric inlet study, the recommended inlet L/D is suggested to be between 0.25 and 0.4. The performance of a candidate short inlet with L/D = 0.25 was assessed using full-annulus unsteady Reynolds-averaged Navier–Stokes (RANS) simulations at critical design and off-design operating conditions. The candidate design maintains the propulsive efficiency of the baseline case and fuel burn benefits are conjectured due to reductions in nacelle weight and drag compared to an aircraft powered by the baseline propulsor.

Addressing the Pushback Time Allocation Problem at Heathrow Airport

This paper considers the problem of allocating pushback times to departing aircraft, specifying the time at which they will be given permission to push back from their allocated stand, start their engines, and commence their taxi to the runway. The aim of this research is to first predict the delay (defined as the waiting time at the stand or runway) for each departure, then to use this to calculate a pushback time such that an appropriate amount of the delay is absorbed at the stand, prior to starting the engines. A two-stage approach is used, where the feasibility of the second stage (pushback time allocation) has to be considered within the first stage (takeoff sequencing). The characteristics of this real-world problem and the differences between it and similar problems are thoroughly discussed, along with a consideration of the important effects of these differences. Differences include a nonlinear objective function with a nonconvex component; the integration of two sequence dependent separation problems; separations that can vary over time; and time-slot extensions. Each of these factors has contributed to the design of the solution algorithm. Results predict significant fuel-burn benefits from absorbing some of the delay as stand hold, as well as delay benefits from indirectly aiding the runway controllers by reducing runway queue sizes. A system for pushback time allocation at London Heathrow has been developed by NATS (formerly National Air Traffic Services) based upon the algorithm described in this paper.

Effects of Advanced Engine Technology on Open Rotor Cycle Selection and Performance

Recent increases in fuel prices and increased focus on aviation's environmental impacts have reignited focus on the open rotor engine concept. This type of architecture was extensively investigated in previous decades but was not pursued through to commercialization due to relatively high noise levels and a sudden, sharp decrease in fuel prices. More recent increases in fuel prices and increased government pressure from taxing carbon-dioxide production mean the open rotor is once again being investigated as a viable concept. Advances in aero-acoustic design tools have allowed industry and academia to re-investigate the open rotor with an increased emphasis on noise reduction while retaining the fuel burn benefits due to the increased propulsive efficiency. Recent research with conceptual level multidisciplinary considerations of the open rotor has been performed (Bellocq et al., 2010, “Advanced Open Rotor Performance Modeling For Multidisciplinary Optimization Assessments,” Paper No. GT2010-2963), but there remains a need for a holistic approach that includes the coupled effects of the engine and airframe on fuel burn, emissions, and noise. Years of research at Georgia Institute of Technology have led to the development of the Environmental Design Space (EDS) (Kirby and Mavris, 2008, “The Environmental Design Space,” Proceedings of the 26th International Congress of the Aeronautical Sciences). EDS serves to capture interdependencies at the conceptual design level of fuel burn, emissions, and noise for conventional and advanced engine and airframe architectures. Recently, leveraging NASA environmentally responsible aviation (ERA) modeling efforts, EDS has been updated to include an open rotor model to capture, in an integrated fashion, the effects of an open rotor on conventional airframe designs. Due to the object oriented nature of EDS, the focus has been on designing modular elements that can be updated as research progresses. A power management scheme has also been developed with the future capability to trade between fuel efficiency and noise using the variable pitch propeller system. Since the original GE open rotor test was performed using a military core, there is interest in seeing the effect of modern core-engine technology on the integrated open rotor performance. This research applies the modular EDS open rotor model in an engine cycle study to investigate the sensitivity of thermal efficiency improvements on open rotor performance, including the effects on weight and vehicle performance. The results are that advances in the core cycle are necessary to enable future bypass ratio growth and the trades between core operating temperatures and size become more significant as bypass ratio continues to increase. A general benefit of a 30% reduction in block fuel is seen on a 737-800 sized aircraft.