A novel full-electric aircraft propulsion based on the DEA compressor

2021 ◽  
pp. 095441002110294
Author(s):  
Babak Aryana
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Sea Level ◽  
Static Condition ◽  
Electric Aircraft ◽  
Specific Thrust ◽  
Transonic Region ◽  
Aircraft Propulsion

This article introduces a novel full-electric aircraft propulsion designed based on the DEA compressor, named DEAThruster that is supplied by a PEMFC. The thruster designed and modeled in this study is a compact DEAThruster to operate in altitudes up to 15,000 m and in the subsonic/transonic region with a specific thrust around 8300 N s/kg. The results show the thruster can satisfy all expectations, and it can generate up to 8300 N s/kg specific thrust for flight conditions encompassing static condition at sea level up to flight Mach number 0.95 in altitude 15,000 m. The DEAThruster can potentially be a practical alternative for gas turbine propulsions in all aspects when all available options for full-electric propulsions are not competitive for conventional aircraft propulsions in performance, size, and weight.

A novel hybrid aircraft propulsion based on the DEA compressor—part B: Performance

2021 ◽  
pp. 095441002110287
Author(s):  
Babak Aryana
Keyword(s):  
Mach Number ◽  
Sea Level ◽  
Static Condition ◽  
Propulsion System ◽  
Axial Turbine ◽  
Hybrid Propulsion ◽  
Pulse Detonation ◽  
Exit Nozzle ◽  
Distributed Propulsion ◽  
Detonation Process

This two-parts article introduces a novel hybrid propulsion system based on the DEA compressor. The system encompasses a Pulse Detonation TurboDEA as the master engine that supplies several full-electric ancillary thrusters called DEAThruster. The system, called the propulsion set, can be categorized as a distributed propulsion system based on the design mission and number of ancillary thrusters. Part B of this article explains the performance sizing of the propulsion set designed in part A. Evaluating the performance of the propulsion, computer programs are written for all major components of the both master engine and ancillary thruster. The intake, compressor, detonation process, diffusers, axial turbine, and exit nozzle are modeled under certain flight conditions, and their performances are revealed and analyzed. The flight conditions are considered from the static condition at the sea level up to flight Mach number 5 at an altitude of 20,000 m. The performance of the propulsion set is also compared with some aircraft propulsions modeled by similar studies in all important aspects.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

System Weight Comparison of Electric Machine Topologies for Electric Aircraft Propulsion

2018 ◽  
Author(s):  
Aaron D. Anderson ◽  
Nathaniel J. Renner ◽  
Yuyao Wang ◽  
Dongsu Lee ◽  
Shivang Agrawal ◽  
...  
Keyword(s):  
Electric Machine ◽  
Electric Aircraft ◽  
Aircraft Propulsion

Method for calculating the thrust and specific thrust of a double-flow gas turbine engine based on the results of tests at the laboratory unit «Flame»

2021 ◽  
pp. 58-64
Author(s):  
S.M. Sergeev ◽  
◽  
V.A. Kudriashov ◽  
N.V. Petrukhin ◽  
◽  
...  
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Similarity Criteria ◽  
Fuel Quality ◽  
Jet Engines ◽  
Laboratory Unit ◽  
Turbine Engine ◽  
Specific Thrust ◽  
The Cost

The main technical characteristics of jet engines depend on the fuel quality: thrust and fuel consumption. As a rule, the comparative assessment of real engines is carried by specific values. Specific thrust is one of the most important parameters of the gas turbine engine (GTE). The larger it is, the smaller the required air flow rate through the engine at a given thrust and therefore its dimensions and mass. To date, a system for evaluating the performance properties of fuels based on qualification methods has been created. However, these methods do not allow calculating the thrust and specific thrust of the engine and potentially assessing the effect of fuels on these characteristics. Therefore, the issues of efficient use of fuels for GTE are solved almost exclusively on the basis of tests at testing units with full-scale engines, which are carried out repeatedly, which leads to a significant increase in the cost of testing. The article proposes a method for calculating the thrust and specific thrust of a double-flow gas turbine engine according to the results of tests at a constant volume laboratory unit of bypass type “Flame”. The method is based on modeling the engine operating conditions using the similarity criteria of the bench reactor and the real engine and allows reducing significantly the material and time costs for testing. The experimental of the combustion characteristics of hydrocarbon fuels and the rated values of their thrust and specific thrust for a double-flow gas turbine engine are presented.

scholarly journals Conceptual Mean-Line Design of Single and Twin-Shaft Oxy-Fuel Gas Turbine in a Semiclosed Oxy-Fuel Combustion Combined Cycle

2013 ◽  
Vol 135 (8) ◽  
Author(s):  
Majed Sammak ◽  
Egill Thorbergsson ◽  
Tomas Grönstedt ◽  
Magnus Genrup
Keyword(s):  
Mach Number ◽  
Gas Turbine ◽  
Mass Flow ◽  
Rotational Speed ◽  
Gas Turbines ◽  
Combined Cycle ◽  
Fuel Gas ◽  
Power Turbine ◽  
Exit Mach Number ◽  
Line Design

The aim of this study was to compare single- and twin-shaft oxy-fuel gas turbines in a semiclosed oxy-fuel combustion combined cycle (SCOC–CC). This paper discussed the turbomachinery preliminary mean-line design of oxy-fuel compressor and turbine. The conceptual turbine design was performed using the axial through-flow code luax-t, developed at Lund University. A tool for conceptual design of axial compressors developed at Chalmers University was used for the design of the compressor. The modeled SCOC–CC gave a net electrical efficiency of 46% and a net power of 106 MW. The production of 95% pure oxygen and the compression of CO2 reduced the gross efficiency of the SCOC–CC by 10 and 2 percentage points, respectively. The designed oxy-fuel gas turbine had a power of 86 MW. The rotational speed of the single-shaft gas turbine was set to 5200 rpm. The designed turbine had four stages, while the compressor had 18 stages. The turbine exit Mach number was calculated to be 0.6 and the calculated value of AN2 was 40 · 106 rpm2m2. The total calculated cooling mass flow was 25% of the compressor mass flow, or 47 kg/s. The relative tip Mach number of the compressor at the first rotor stage was 1.15. The rotational speed of the twin-shaft gas generator was set to 7200 rpm, while that of the power turbine was set to 4800 rpm. A twin-shaft turbine was designed with five turbine stages to maintain the exit Mach number around 0.5. The twin-shaft turbine required a lower exit Mach number to maintain reasonable diffuser performance. The compressor turbine was designed with two stages while the power turbine had three stages. The study showed that a four-stage twin-shaft turbine produced a high exit Mach number. The calculated value of AN2 was 38 · 106 rpm2m2. The total calculated cooling mass flow was 23% of the compressor mass flow, or 44 kg/s. The compressor was designed with 14 stages. The preliminary design parameters of the turbine and compressor were within established industrial ranges. From the results of this study, it was concluded that both single- and twin-shaft oxy-fuel gas turbines have advantages. The choice of a twin-shaft gas turbine can be motivated by the smaller compressor size and the advantage of greater flexibility in operation, mainly in the off-design mode. However, the advantages of a twin-shaft design must be weighed against the inherent simplicity and low cost of the simple single-shaft design.

Numerical Modeling of Heat Transfer and Pressure Losses for an Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2009 ◽  
Vol 1 (2) ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. Computational fluid dynamics (CFD) predictions of blade tip heat transfer are compared with test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; a flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25%, 2.0%, and 2.75% of the blade span. The tip heat transfer results of the numerical models agree well with data. For the case in which side-by-side comparison with test measurements in the open literature is possible, the magnitude of the heat transfer coefficient in the “sweet spot” matches data exactly and shows 20–50% better agreement with experiment than prior CFD predictions of this same case.

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