Panagiotis Laskaridis

Methodology to assess the performance of an aircraft concept with distributed propulsion and boundary layer ingestion using a parametric approach

The performance benefits of boundary layer ingestion in aircraft with distributed propulsion have been extensively studied in the past. These studies have indicated that propulsion system integration issues such as distortion and intake pressure losses could mitigate the expected benefits. This paper introduces and develops a methodology that enables the assessment of different propulsion system designs, which are optimized to be less sensitive to the effects of the aforementioned issues. The study models the propulsor array and main engine performance at design point using a parametric approach, and further at component level, the study focuses on identifying optimum propulsor configurations, in terms of propulsor pressure ratio and BL capture sheet height. At a system level, the study assesses the effects of splitting the thrust between the propulsor array and main engines. The figure of merit used in the optimization is the TSFC. The suitability of the concepts is further assessed using performance predictions for HTS electrical motors. For the purpose of this study, the NASA N3-X aircraft concept is selected as baseline configuration, where the different propulsion designs are tested. As the study focuses on performance assessment of the propulsion system, sizing implication issues and aircraft performance installations effects have not been included in the analysis. The results from the parametric analysis corroborated previous studies regarding the high sensitivity of the propulsion system performance to intake losses and BL inlet conditions. As the study found low-power consumption configurations at these operating conditions, this may be considered as a major issue. The system analysis from the study indicated that splitting the thrust between propulsors and main engines results in improved system efficiency with beneficial effects in fuel savings. When a 2% increase in intake pressure losses and a similar reduction in fan efficiency were assumed due to boundary layer ingestion, the study found an optimum configuration with 65% of thrust delivered by the propulsor array. To summarize, the present work built on past research further contributes to the field through the inclusion of the thrust split as a key variable in the propulsion system design. The thrust split, when introduced, enabled reduction of the detrimental effects of intake losses on the overall system performance. Additionally, as it reduces the power required for the propulsor array, it is expected to reduce the operating power of HTS and cooling systems and therefore improve the effectiveness of the concept.

Performance characteristics and optimisation of a geared intercooled reversed flow core engine

Intercooled turbofan cycles allow higher overall pressure ratios to be reached, which gives rise to improved thermal efficiency. In addition, intercooling allows for the size, weight and exhaust jet velocity of the core to be reduced. For an optimum jet velocity ratio and fixed thrust, the fan pressure ratio and specific thrust are also reduced, which benefits propulsive efficiency. A new intercooled core concept is proposed in this paper, which promises to alleviate limitations identified in previous intercooled turbofan designs. This concept facilitates the installation of the intercooler and reduces core losses at high overall pressure ratios. This engine concept takes advantage of intercooling and the arrangement of the high pressure spool to reach and exceed overall pressure ratios of 80. In addition, given the reduction in core size, bypass ratios beyond 14 have been considered. In order to identify efficiency gains and performance characteristics which are due to the novel arrangement alone, the geared intercooled reversed flow core engine has been compared with a geared intercooled engine with a more conventional core. Finally an optimisation exercise has been carried out to identify the best configuration for both the geared intercooled reversed flow core concept and the conventional core concept. In this paper, it is demonstrated that the geared intercooled reversed flow core concept allows for a 2.3% reduction in block fuel burn. The reductions are due to the improved core efficiency, higher overall pressure ratio as well as efficiency gains from the use of a mixed exhaust. The sensitivity analysis shows that the improvements are highly dependent on pressure losses in the core and bypass stream and that careful design of the mixer chutes and intercooler headers to achieve low losses is essential if the concept gains are to be realised.

Toward a Scalable Propeller Performance Map

Propeller performance is traditionally represented by a performance map that gives propeller efficiency as a function of the flight Mach number, the power coefficient CP, and the advance ratio J. This work aims to demonstrate how this map changes when the design CP and J change and to propose a novel map format that is able to capture the performance of different propeller designs. For this purpose, the propeller performance is simulated using a propeller lifting-line method validated for the SR3 propfan. Subsequently, the propeller model is used within a sequential quadratic programming framework to optimize the blade twist and chord distribution for different sets of design CP and J. A complete propeller performance map is then generated for each one of the optimized designs. The results demonstrate that all the investigated propellers can be modeled by a common map, which determines separately the ideal efficiency and the viscous losses. The ideal efficiency is given in the traditional format of ηi=f(C...

Severity estimation and effect of operational parameters for civil aircraft jet engines

Engine maintenance costs for civil aircraft largely depend on the life consumption of critical parts like high-pressure turbine blade and disk. Severity (defined as a measure of relative damage) can be used to characterize the degree of life consumption for the parts subjected to low cycle fatigue, creep, and oxidation as predominant life-limiting modes under normal engine operation. In this context, a numerical tool has been developed to study the impact of engine-aircraft operational parameters on engine severity. This is accomplished by combining aircraft performance, gas turbine performance, component sizing, heat transfer calculation, thermal analysis, structural analysis, life and severity estimation for the high-pressure turbine, and the disk components of jet engines. Operational parameters such as take-off (TO) derate, climb derate, outside air temperature, and airport altitude have been studied for lower thrust and larger thrust engines. The observations from the study show that TO derate and outside air temperature are dominant parameters in terms of engine severity. The use of climb derate also lowers the severity and is more beneficial for larger thrust engines, while airport altitude effects could be minimized by the use of suitable TO derate. The severity characteristics estimated through the numerical tool can provide an enhanced perspective on aircraft engine maintenance.

Ejector Pump Theory Applied to Gas Turbine Engine Performance inside Indoor Sea-Level Test Cell-Analytical and CFD Study

The paper deals with the interaction between the engine exhaust jets and the detuner of indoor gas turbine test facilities. This is part of an on-going research programme at Cranfield University looking at sea-level gas turbine testing from several aspects. In the case of sea level indoor testing several flow phenomena affecting both the engine performance and the thrust measurement arise inside the cell. Most of these phenomena are strictly related to the amount of secondary flow entering the cell which depends mainly on the interaction between the engine exhaust jets and the detuner. Following the above considerations this paper, beside a review of the main parameters that affect the performance of an ejector pump, concentrates on the development and application of an analytical ejector pump model. A complete description of the analytical model is presented including results obtained using genuine test data. The study covers the derivation of the Engine-Detuner ejector characteristic derived by using commercial CFD packages. Several computational models have been generated to analyze the influences of different Engine-Cell configurations on the performance characteristic of the ejector pump taking place inside indoor test cell. A comparison between analytical and CFD results is also presented.

Concept description and assessment of the main features of a geared intercooled reversed flow core engine

Intercooled turbofan cycles allow higher overall pressure ratios to be reached which gives rise to improved thermal efficiency. Intercooling also allows core mass flow rate to be reduced which facilitates higher bypass ratios. A new intercooled core concept is proposed in this paper which promises to alleviate limitations identified with previous intercooled turbofan designs. Specifically, these limitations are related to core losses at high overall pressure ratios as well as difficulties with the installation of the intercooler. The main features of the geared intercooled reversed flow core engine are described. These include an intercooled core, a rear-mounted high-pressure spool fitted rearwards of the low-pressure spool as opposed to concentrically as well as a mixed exhaust. In these studies, the geared intercooled reversed flow core engine has been compared with a geared intercooled straight flow core engine with a more conventional core layout. This paper compares the mechanical design of the high-pressure spools and shows how different high-pressure compressor and high-pressure turbine blade heights can affect over-tip leakage losses. In the reversed configuration, the reduction in high-pressure spool mean diameter allows for taller high-pressure compressor and turbine blades to be adopted which reduces over-tip leakage losses. The implication of intercooler sizing and configuration, including the impact of different matrix dimensions, is assessed for the reversed configuration. It was found that a 1-pass intercooler would be more compact although a 2-pass would be less challenging to manufacture. The mixer performance of the reversed configuration was evaluated at different levels of mixing effectiveness. This paper shows that the optimum ratio of total pressure in the mixing plane for the reversed flow core configuration is about 1.02 for a mixing effectiveness of 80%. Lower mixing effectiveness would result in a higher optimum ratio of total pressure in the mixing plane and fan pressure ratio.

A preliminary method to estimate impacts of inlet flow distortion on boundary layer ingesting propulsion system design point performance

Boundary layer ingestion by fans in propulsion system improves the propulsive efficiency. However, inlet flow distortion will dramatically eliminate these benefits. This paper puts forward a method to deal with inlet flow distortions and examines their impacts on turbofan performance at engine design point. The method models both radial and circumferential distortion and their impacts separately. Firstly, a distorted fan map is calculated by parallel compressor method. Then, the new map is utilised to find the fan exit flow conditions by parallel stream method. Finally, we assume that all the flows mixed well before entering the nozzle without any pressure losses. At all examined fan pressure ratios, boundary layer ingesting improved fuel consumption. However, the benefits reduced by the new method are lower than previous predictions without considering intake distortion. If the fan pressure loss and efficiency drop due to inlet distortion are too high, boundary layer ingestion should not be used with a traditional fan design. Large boundary layer ingestion for future propulsion system should consider new fan blade design.

Industrial Gas Turbine Performance: Compressor Fouling and On-Line Washing

Industrial gas turbines are susceptible to compressor fouling, which is the deposition and accretion of airborne particles or contaminants on the compressor blades. This paper demonstrates the blade aerodynamic effects of fouling through experimental compressor cascade tests and the accompanied engine performance degradation using turbomatch, an in-house gas turbine performance software. Similarly, on-line compressor washing is implemented taking into account typical operating conditions comparable with industry high pressure washing. The fouling study shows the changes in the individual stage maps of the compressor in this condition, the impact of degradation during part-load, influence of control variables, and the identification of key parameters to ascertain fouling levels. Applying demineralized water for 10 min, with a liquid-to-air ratio of 0.2%, the aerodynamic performance of the blade is shown to improve, however most of the cleaning effect occurred in the first 5 min. The most effectively washed part of the blade was the pressure side, in which most of the particles deposited during the accelerated fouling. The simulation of fouled and washed engine conditions indicates 30% recovery of the lost power due to washing.

Turboelectric Distributed Propulsion System Modelling for Hybrid-Wing- Body Aircraft

(ABSTRACT) In 2005, NASA released plans of next generation commercial airplane for 2030, with a crossdisciplinary effort on: reduced fuel consumption, aviation reliability, fundamental noise reduction and shorter take-off length. Meeting these requirements will need a fundamental shift in aircraft and engine design. Turboelectric distributed propulsion system was chosen to achieve these targets. Different from traditional turbofan, distributed propulsion system employs a large number of fans embedded on upper surface of the airframe and two turbogenerators at wing tip. This novel configuration benefits from boundary layer ingestion and distributed fans to achieve higher bypass ratio but lower fuel burn. The N3-X hybrid-wing-body is used as a baseline aircraft for the study. This paper gives basic simulation methods, as well as computational models for turboelectric distributed propulsion system. Initially, a boundary layer ingesting model has been built from computational results for embedded propulsor at different inlet conditions. In a further step, a weight estimation model of propulsors was concluded to estimate propulsors’ weight and size. Then, thermal cycle model was built to calculate engine’s performance at both design point and off design conditions. Finally, effects of boundary layer ingestion on the propulsion system were examined. The boundary layer ingesting model showed mass-average inlet pressure and Mach number are function of flight Mach number and fan inlet mass flow, on the N3-X airframe. The weight estimation model shows the overall system weight decreased with increased number of propulsors, which also caused total inlet width of propulsors increasing. So for a given total inlet width, the propulsor should be used as many as possible to reduce weight. Thermal cycle results show that fan shaft speed should be chosen as high as possible before reaching the fan tip speed limitation, and fan pressure ratio (FPR) between 1.3 and 1.35 yields minimum thrust specific fuel consumption (TSFC) at the aerodynamic design point. A fan pressure ratio of 1.3 is chosen for its potential effects on noise control. In the end, a turboelectric distributed engine was simulated to satisfy NASA N+3 subsonic commercial airplane goals.

Effects of Off-takes for Aircraft Secondary-Power Systems on Jet Engine Efficiency

Power and bleed-air engine offtakes have a detrimental effect on the efficiency of the engine. This paper presents a thermodynamic analysis that relates the offtakes specific fuel-consumption penalties to the design parameters of the engine. A set of equations is derived in order to determine the fuel-consumption penalties of known power and bleedair offtakes. Their predictions are compared against numerical results performedwith an in-house engine simulation tool. The equations include design parameters that are known in the early preliminary design stages and can be a useful tool for aircraft, engine, and secondary-systems designers. Furthermore, the analytical expressions offer physical insight on the parameters that drive the magnitude of the specific fuel-consumption penalties. It has been found that the main factor driving the magnitude of the penalties is the size of the offtakes relative to the core power. This leads to the conclusion that future engineswith a lower specific thrust factorST=V0, amore efficient fan and lowpressure turbine, and a higher power factor Ppo= T V0 will face increased offtake penalties.