Potential of Micro Turbine Based Propulsion Systems for Civil UAVs: A Case Study

2016 ◽  
Author(s):  
A. Marcellan ◽  
W. P. J. Visser ◽  
P. Colonna
Keyword(s):  
Scale Effects ◽  
Pressure Ratio ◽  
Performance Model ◽  
Inlet Temperature ◽  
Propulsion Systems ◽  
University Of Technology ◽  
Turboprop Engine ◽  
Component Efficiency ◽  
Micro Turbine ◽  
Technological Improvements

There is a high potential for civil applications of Unmanned Aerial Vehicles (UAV) in areas such as goods transport, telecommunication, remote monitoring and sensing, surveillance, search and rescue, and disaster management. Developments in areas such as telecommunication, control and information technology offer opportunities for long range remotely or automatically piloted missions. This requires efficient and light-weight small propulsion systems. The potential of turboprop propulsion for civil UAVs using micro turbine technology has been explored and compared with existing concepts, such as piston engine driven propellers. Different propulsion concepts have been analyzed and the application areas where advanced turboprops would be superior to other systems such as reciprocating engines and electric motors, identified. However, turboprop engines of the small power capacity required for the aircraft concepts and missions considered are not currently available with competitive performance. A conceptual design study of a micro turboprop engine has been performed by downscaling an existing reference engine. Scale effects on efficiency have been taken into account, as well as effects of technological progress. Engine cycle optimization has been carried out and the effects of turbine inlet temperature, compressor pressure ratio, engine size, and component efficiency have been investigated. An aerodynamic and flight performance model of a baseline UAV has been developed in order to predict mission performance. This model has been coupled to a turboprop model to evaluate system performance with different engine configurations for the selected mission. The outcome of the study provides information about the technological improvements in terms of cycle efficiency required to make the micro-turboprop a competitive solution. The Propulsion and Power group of Delft University of Technology will pursue these R&D goals in an attempt to contribute to the development of civil UAV technology.

Thermo-Economic Optimization in the Design of Small-Scale and Residential Cogeneration Systems

2009 ◽  
Author(s):  
C. P. Lea˜o ◽  
S. F. C. F. Teixeira ◽  
A. M. Silva ◽  
M. L. Nunes ◽  
L. A. S. B. Martins
Keyword(s):  
Gas Turbine ◽  
Urban Areas ◽  
Pressure Ratio ◽  
Optimization Method ◽  
Turbine Engines ◽  
Small Scale ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Economic Optimization ◽  
Micro Turbine

In recent years, gas-turbine engines have undergone major improvements both in efficiency and cost reductions. Several inexpensive models are available in the range of 30 to 250 kWe, with electrical efficiencies already approaching 30%, due to the use of a basic air-compressor associated to an internal air pre-heater. Gas-turbine engines offer significant advantages over Diesel or IC engines, particularly when Natural Gas (NG) is used as fuel. With the current market trends toward Distributed Generation (DG) and the increased substitution of boilers by NG-fuelled cogeneration installations for CO2 emissions reduction, small-scale gas turbine units can be the ideal solution for energy systems located in urban areas. A numerical optimization method was applied to a small-scale unit delivering 100 kW of power and 0.86 kg/s of water, heated from 318 to 353K. In this academic study, the unit is based on a micro gas-turbine and includes an internal pre-heater, typical of these low pressure-ratio turbines, and an external heat recovery system. The problem was formulated as a non-linear optimisation model with the minimisation of costs subject to the physical and thermodynamic constraints. Despite difficulties in obtaining data for some of the components cost-equations, the preliminary results indicate that the optimal compressor pressure ratio is about half of the usual values found in large installations, but higher than those of the currently available micro-turbine models, while the turbine inlet temperature remains virtually unchanged.

Performance Enhancement of Advanced Combined Cycles

2003 ◽  
Author(s):  
Sanjay ◽  
Onkar Singh ◽  
B. N. Prasad
Keyword(s):  
Gas Turbine ◽  
Steam Generator ◽  
Heat Recovery ◽  
Performance Enhancement ◽  
Pressure Ratio ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Heat Recovery Steam Generator ◽  
Component Efficiency ◽  
Plant Efficiency

This paper reports on the development requirements of gas/steam combined cycle with an aim to achieve plant efficiency greater than 62% through various development possibilities in gas turbine and steam turbine cycle by taking a reference combined cycle configuration (MS9001H gas turbine and three pressure heat recovery steam generator with reheat). The innovative development possibilities include the advanced inlet design to reduce pressure loss, the increase in turbine inlet temperature, use of advanced turbine blade material, increased component efficiency, improved turbine cooling technologies along with better cooling medium, incorporating intercooling, reheat and regeneration either separately or in combination with simple gas turbine cycle using higher compressor pressure ratio, better utilization of heat recovery steam generator, minimum stack temperature, single shaft system configuration, etc. Based on the quantification of each development item, if incorporated in reference cycle, it has been estimated that the combined cycle as the potential to achieve the plant efficiency in excess of 63%.

scholarly journals Assessment of Surface Roughness Effects on Micro Axial Turbines

2020 ◽  
Author(s):  
Abdelaziz A. A. Gamil ◽  
Theoklis Nikolaidis ◽  
Joao A. Teixeira ◽  
S. H. Madani ◽  
Ali Izadi
Keyword(s):  
Heat Transfer ◽  
Surface Roughness ◽  
Gas Turbine ◽  
Pressure Ratio ◽  
Careful Consideration ◽  
Inlet Temperature ◽  
Total Efficiency ◽  
Roughness Effects ◽  
Micro Gas Turbine ◽  
Micro Turbine

Abstract Surface roughness significantly affects the aerodynamics and heat transfer within micro-scale turbine stages. It results in a considerable increment in the blade profile loss and leads consequently to sizeable performance reductions. The provision of low roughness surfaces in micro gas turbine stages presents challenges on account of the small (mm scale) sizes, manufacturing complexity and associated costs. The axial turbine investigated in this study is fitted to Samad Power’s TwinGen domestic micro combined heat and power unit. The micro gas turbine has a compressor pressure ratio of 3, 1200K turbine inlet temperature and a rotational speed of 170,000 rpm. This paper presents a numerical assessment of the effects of varying the surface roughness on the performance and heat transfer of the micro turbine. The surface roughness was uniformly distributed on the NGV and rotor blades. The results showed that increasing the surface roughness from 3 microns to 6, 20, and 100 microns resulted in a reduction in stage total efficiency of 0.8%, 4% and 12% respectively as well as a comparable decrease in output power (0.7%, 3.6%, and 11% respectively). The turbine temperature was also observed to be very sensitive to surface roughness and a temperature increase of some 5% at the rotor hub and over 4% increment in the blade tip surface was observed for 100 microns when compared to the 3 microns surface roughness case. The findings of this paper highlight the adverse effects of the surface roughness on the micro-turbine performance and temperature distribution as well as the importance of careful consideration of wall roughness during the design and manufacturing stages.

Performance Simulation of Turboprop Engine Using SIMULINK® Model

2002 ◽  
Author(s):  
Changduk Kong ◽  
Hongsuk Roh
Keyword(s):  
Steady State ◽  
Sea Level ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Performance Parameters ◽  
Performance Simulation ◽  
Simulink Model ◽  
First Case ◽  
Turboprop Engine

A performance simulation model of a turboprop engine, the PT6A-62, which is the power plant of KT-1, was developed to predict the steady-state behaviors using the SIMULINK® model. The SIMULINK model consists of subsystems to represent engine components such as intake, compressor, combustor, compressor turbine, power turbine and exhaust nozzle. For validation, performance parameters calculated using the SIMULINK model were compared with the results using GASTURB model. The steady-state performance analysis using the developed SIMULINK model was performed. Performance parameters, such as the mass flow rate, the compressor pressure ratio, the fuel flow rate, the specific fuel consumption ratio and the turbine inlet temperature, were conducted to evaluate validity of the SIMULINK model at various cases. The first case was the uninstalled condition at various altitudes from sea level to 9144m (30000ft) with fixed M.N. = 0. And the second case was the installed condition at various altitudes from sea level to 7620m (25000ft) with fixed M.N. = 0. The third case was the installed condition at altitudes of 1524m (5000ft) and 3048m (1000ft) and at the M.N. = 0.1, 0.2 and 0.3 in ECS operation ECS. In this investigation, it was confirmed that the results using the SIMULINK model were well agreed with the results using the GASTURB model within maximum 6.5%.

scholarly journals PZL-10 Turboshaft Engine–System Design Review

2019 ◽  
Vol 26 (1) ◽  
pp. 23-29
Author(s):  
Michal Czarnecki ◽  
John Olsen ◽  
Ruixian Ma
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Performance Model ◽  
Inlet Temperature ◽  
Gas Generator ◽  
Design Bureau ◽  
Turbine Engine ◽  
Engine Design ◽  
Turboshaft Engine

Abstract The PZL – 10-turboshaft gas turbine engine is straight derivative of GTD-10 turboshaft design by OKMB (Omsk Engine Design Bureau). Prototype engine first run take place in 1968. Selected engine is interested platform to modify due gas generator layout 6A+R-2, which is modern. For example axial compressor design from successful Klimov designs TB2-117 (10A-2-2) or TB3-117 (12A-2-2) become obsolete in favour to TB7-117B (5A+R-2-2). In comparison to competitive engines: Klimov TB3-117 (1974 – Mi-14/17/24), General Electric T-700 (1970 – UH60/AH64), Turbomeca Makila (1976 – II225M) the PZL-10 engine design is limited by asymmetric power turbine design layout. This layout is common to early turboshaft design such as Soloview D-25V (Mil-6 power plant). Presented article review base engine configuration (6A+R+2+1). Proposed modifications are divided into different variants in terms of design complexity. Simplest variant is limited to increase turbine inlet temperature (TIT) by safe margin. Advanced configuration replace engine layout to 5A+R+2-2 and increase engine compressor pressure ratio to 9.4:1. Upgraded configuration after modification offers increase of generated power by 28% and SFC reduction by 9% – validated by gas turbine performance model. Design proposal corresponds to a major trend of increasing available power for helicopter engines – Mi-8T to Mi-8MT – 46%, H225M – Makila 1A to 1A2 — 9%), Makila 1A2 to Makila 2-25%.

Thermodynanic Analysis of a Combined Micro Turbine With a Micro ORC

2008 ◽  
Author(s):  
Mortaza Yari
Keyword(s):  
Pressure Ratio ◽  
Organic Rankine Cycle ◽  
Internal Heat ◽  
Rankine Cycle ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Evaporator Temperature ◽  
Combined Cycles ◽  
Micro Turbine ◽  
Organic Fluids

In the last years, a big effort has been undergone to improve micro turbines thermal efficiency, actually rated at about 30%. A value of 40% is often regarded as a possible target. Such a result could be achieved implementing more complicated thermodynamic cycles, like combined cycles. This paper deals with the hypothesis of bottoming a low pressure ratio, recuperated gas cycle, typically realized in actual micro turbines, with an Organic Rankine Cycle (ORC) with internal heat exchanger (IHE), obtaining a micro-combined-cycle. The results are presented and the influence of the several parameters: Turbine inlet temperature of the micro turbine, compressor pressure ratio, evaporation temperature and evaporator temperature difference are discussed. Both simple ORC and ORC with IHE bottom cycle options are discussed. The dry organic fluids in this study are Isopentane, n-Pentane, n-Heptane, n-Octane, n-Hexane, R113, R123 and Toluene.

Propulsive Performance of Ideal Detonation Turbine Based Combined Cycle Engine

2012 ◽  
Vol 134 (8) ◽  
Author(s):  
Hua Qiu ◽  
Cha Xiong ◽  
Chuan-jun Yan ◽  
Wei Fan
Keyword(s):  
Pressure Ratio ◽  
Area Ratio ◽  
Combined Cycle ◽  
Performance Model ◽  
Propulsion System ◽  
Pulse Detonation Engine ◽  
Inlet Temperature ◽  
Pulse Detonation ◽  
Propulsive Performance ◽  
Detonation Engine

A novel two-mode propulsion system based on detonation combustion, known as a detonation turbine based combined cycle engine (DTBCC), was proposed and thermodynamically analyzed for potential application to aircrafts whose flight Mach number is from 0 to 5. The obvious advantage of the two-mode system is that both modes share the same multidetonation chambers. The quasi-stable total temperature and total pressure for inlet conditions of the turbine could be realized in this hybrid pulse detonation engine. A key parameter (drive area ratio) was defined as the ratio of the outflow area at the head to the cross-sectional area of the detonation chamber. The calculated results showed that the increase of the drive area ratio led to the increase in the mass flow entering the turbine; however, this led to the decrease of the total inlet temperature, the total inlet pressure, and the expansion-pressure ratio of the turbine. Compared with an ideal turbojet engine, the inlet temperature of the turbine in a preturbine hybrid pulse detonation engine with a drive area ratio of 1 was 80 K lower than the former under the same pressure ratio and the same fuel-air ratio. In other words, the increase of the drive area ratio may improve the performance of this hybrid pulse detonation engine. Variation of the pressure ratio was adapted to varied flight Mach numbers by a change of the drive area ratio, which induced the enlargement of the operating range. Finally, a performance model was established to research the components’ characteristics and the propulsive performance of the engine. Preliminary performance estimates suggested that thrust and specific fuel consumption of the two-mode propulsion system were superior to the existing turbine based combined cycle designs.

scholarly journals Feasibility of a Helium Closed-Cycle Gas Turbine for UAV Propulsion

2020 ◽  
Vol 11 (1) ◽  
pp. 28
Author(s):  
Emmanuel O. Osigwe ◽  
Arnold Gad-Briggs ◽  
Theoklis Nikolaidis
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Low Noise ◽  
Closed Cycle ◽  
Propulsion Systems ◽  
Component Design ◽  
Readiness Level ◽  
Aerial Vehicle ◽  
Turbine Design

When selecting a design for an unmanned aerial vehicle, the choice of the propulsion system is vital in terms of mission requirements, sustainability, usability, noise, controllability, reliability and technology readiness level (TRL). This study analyses the various propulsion systems used in unmanned aerial vehicles (UAVs), paying particular focus on the closed-cycle propulsion systems. The study also investigates the feasibility of using helium closed-cycle gas turbines for UAV propulsion, highlighting the merits and demerits of helium closed-cycle gas turbines. Some of the advantages mentioned include high payload, low noise and high altitude mission ability; while the major drawbacks include a heat sink, nuclear hazard radiation and the shield weight. A preliminary assessment of the cycle showed that a pressure ratio of 4, turbine entry temperature (TET) of 800 °C and mass flow of 50 kg/s could be used to achieve a lightweight helium closed-cycle gas turbine design for UAV mission considering component design constraints.

The Effects of Ambient Conditions on Gas Turbine Emissions—Generalized Correction Factors

1978 ◽  
Vol 100 (4) ◽  
pp. 640-646 ◽  
Author(s):  
P. Donovan ◽  
T. Cackette
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Turbine Engines ◽  
Inlet Temperature ◽  
Gas Turbine Engines ◽  
Correction Factors ◽  
Oxides Of Nitrogen ◽  
Ambient Conditions ◽  
Nitrogen Emission ◽  
Wide Range

A set of factors which reduces the variability due to ambient conditions of the hydrocarbon, carbon monoxide, and oxides of nitrogen emission indices has been developed. These factors can be used to correct an emission index to reference day ambient conditions. The correction factors, which vary with engine rated pressure ratio for NOx and idle pressure ratio for HC and CO, can be applied to a wide range of current technology gas turbine engines. The factors are a function of only the combustor inlet temperature and ambient humidity.

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