Effect of Pressure Ratio on Bending Mode Flutter in a Transonic Linear Cascade

2017 ◽  
Author(s):  
Prahallada Jutur ◽  
Raghuraman N. Govardhan
Keyword(s):  
Structural Damage ◽  
Pressure Ratio ◽  
Aircraft Engine ◽  
Cam Mechanism ◽  
High Frequency Oscillations ◽  
Reduced Frequency ◽  
Bending Mode ◽  
Flow Features ◽  
Engine Components ◽  
Transonic Cascade

Vibration related issues such as flutter have always been a cause of concern for aircraft engine designers. They not only incur unwarranted cost and time overruns, but also significantly compromise performance and can cause structural damage. This phenomenon has become more relevant for the modern aircraft engines, which employ relatively thin, long blade rows to satisfy ever growing demand for a powerful yet compact engine. The tip sections of such blade rows operate with supersonic relative velocity, where prediction of flutter can get challenging due to unsteady flow features like oscillating shocks and their interaction with the blade motion. Linear cascades that represent a specific radial location of the rotor have proven to be a reliable tool for flutter studies. To facilitate flutter experiments at flow Mach numbers realistic to the aircraft engine components, a transonic cascade facility operating at a Mach Number (M) of 1.3 with the ability to oscillate the central blade in the cascade has been developed. The cascade consists of 5 blades and two false blades of which the central blade is oscillated in heave, which represents the bending mode of the rotor. The typical reduced frequencies associated with this kind of flutter in practice (k ∼ 0.1) correspond to a high dimensional frequency of 200 Hz for the present case. A barrel cam mechanism is used to provide such high frequency oscillations. The parameters varied in the present study include the reduced frequency (k) and the static pressure ratio (SPR) across the cascade, which is varied with the help of tailboard and flap arrangement located at the back end of the cascade. Three SPR cases of 1.05, 1.25, and 1.35 are considered and at each of these pressure ratio cases, the reduced frequency is varied. The unsteady loads are measured on the oscillating central blade during the oscillation cycle to quantify the energy transfer from flow to blade and shadowgraphy is used to visualize the shocks. The results from these experiments indicate flutter at lower k values for all the SPR cases tested, while the higher k values are damped. The magnitude of excitation or damping at any particular frequency is also observed to increase with increasing SPR.

Emission Reduction of Fuel Staged Aircraft Engine Combustor Using an Additional Premixed Fuel Nozzle

2012 ◽  
Author(s):  
Takeshi Yamamoto ◽  
Kazuo Shimodaira ◽  
Seiji Yoshida ◽  
Yoji Kurosawa
Keyword(s):  
Pressure Ratio ◽  
Aircraft Engine ◽  
Test Results ◽  
Drastic Reduction ◽  
Fuel Nozzle ◽  
Conducting Research ◽  
Japan Aerospace Exploration Agency ◽  
Co Emission ◽  
Fuel Burner ◽  
Additional Fuel

The Japan Aerospace Exploration Agency (JAXA) is conducting research and development on aircraft engine technologies to reduce environmental impact for the TechCLEAN project. As a part of the project, combustion technologies have been developed with an aggressive target that is an 80% reduction over the NOx threshold of the ICAO CAEP/4 standard. A staged fuel nozzle with a pilot mixer and a main mixer was developed and tested using a single-sector combustor under the target engine’s LTO cycle conditions with a rated output of 40 kN and an overall pressure ratio of 25.8. The test results showed a 77% reduction over the CAEP/4 NOx standard. A reduction in smoke was found under a higher thrust condition than the 30% MTO condition, and a reduction in CO emission was found under a lower thrust condition than the 85% MTO condition. In the present study, an additional fuel burner was designed and tested with the staged fuel nozzle in a single-sector combustor to control emissions. The test results show that the combustor enables an 82% reduction in NOx emissions relative to the ICAO CAEP/4 standard and a drastic reduction in smoke and CO emissions.

scholarly journals Emission Reduction of Fuel-Staged Aircraft Engine Combustor Using an Additional Premixed Fuel Nozzle

2013 ◽  
Vol 135 (3) ◽  
Author(s):  
Takeshi Yamamoto ◽  
Kazuo Shimodaira ◽  
Seiji Yoshida ◽  
Yoji Kurosawa
Keyword(s):  
Technology Development ◽  
Pressure Ratio ◽  
Aircraft Engine ◽  
Development Project ◽  
Civil Aviation ◽  
Test Results ◽  
Technology Development Project ◽  
Fuel Nozzle ◽  
Conducting Research ◽  
Japan Aerospace Exploration Agency

The Japan Aerospace Exploration Agency (JAXA) is conducting research and development on aircraft engine technologies to reduce environmental impact for the Technology Development Project for Clean Engines (TechCLEAN). As a part of the project, combustion technologies have been developed with an aggressive target that is an 80% reduction over the NOx threshold of the International Civil Aviation Organization (ICAO) Committee on Aviation Environmental Protection (CAEP)/4 standard. A staged fuel nozzle with a pilot mixer and a main mixer was developed and tested using a single-sector combustor under the target engine's landing and takeoff (LTO) cycle conditions with a rated output of 40 kN and an overall pressure ratio of 25.8. The test results showed a 77% reduction over the CAEP/4 NOx standard. However, the reduction in smoke at thrust conditions higher than the 30% MTO condition and of CO emission at thrust conditions lower than the 85% MTO condition are necessary. In the present study, an additional fuel burner was designed and tested with the staged fuel nozzle in a single-sector combustor to control emissions. The test results show that the combustor enables an 82% reduction in NOx emissions relative to the ICAO CAEP/4 standard and a drastic reduction in smoke and CO emissions.

An Investigation of Flow Characteristics and Parameter Effects for a New Concept of Hybrid SVC Nozzle

2018 ◽  
Author(s):  
F. Song ◽  
J. W. Shi ◽  
L. Zhou ◽  
Z. X. Wang ◽  
X. B. Zhang
Keyword(s):  
Static Pressure ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Exhaust System ◽  
Aircraft Engine ◽  
Thrust Coefficient ◽  
Thrust Vectoring ◽  
Static Pressure Distribution ◽  
Thrust Vector ◽  
Vector Angle

Lighter weight, simpler structure, higher vectoring efficiency and faster vector response are recent trends in development of aircraft engine exhaust system. To meet these new challenges, a concept of hybrid SVC nozzle was proposed in this work to achieve thrust vectoring by adopting a rotatable valve and by introducing a secondary flow injection. In this paper, we numerically investigated the flow mechanism of the hybrid SVC nozzle. Nozzle performance (e.g. the thrust vector angle and the thrust coefficient) was studied with consideration of the influence of aerodynamic and geometric parameters, such as the nozzle pressure ratio (NPR), the secondary pressure ratio (SPR) and the deflection angle of the rotatable valve (θ). The numerical results indicate that the introductions of the rotatable valve and the secondary injection induce an asymmetrically distributed static pressure to nozzle internal walls. Such static pressure distribution generates a side force on the primary flow, thereby achieving thrust vectoring. Both the thrust vector angle and vectoring efficiency can be enhanced by reducing NPR or by increasing θ. A maximum vector angle of 16.7 ° is attained while NPR is 3 and the corresponding vectoring efficiency is 6.33 °/%. The vector angle first increases and then decreases along with the elevation of SPR, and there exists an optimum value of SPR for maximum thrust vector angle. The effects of θ and SPR on the thrust coefficient were found to be insignificant. The rotatable valve can be utilized to improve vectoring efficiency and to control the vector angle as expected.

scholarly journals Repair of Precision Castings Made of the Inconel 713C Alloy

2017 ◽  
Vol 17 (3) ◽  
pp. 210-216
Author(s):  
K. Łyczkowska ◽  
J. Adamiec
Keyword(s):  
Model Test ◽  
Aircraft Engine ◽  
Casting Defects ◽  
Exhaust Gas ◽  
Plasma Arc ◽  
Nickel Based Superalloy ◽  
Melted Zone ◽  
Base Materials ◽  
Inconel 713C ◽  
Engine Components

Abstract Inconel 713C precision castings are used as aircraft engine components exposed to high temperatures and the aggressive exhaust gas environment. Industrial experience has shown that precision-cast components of such complexity contain casting defects like microshrinkage, porosity, and cracks. This necessitates the development of repair technologies for castings of this type. This paper presents the results of metallographic examinations of melted areas and clad welds on the Inconel 713C nickel-based superalloy, made by TIG, plasma arc, and laser. The cladding process was carried out on model test plates in order to determine the technological and material-related problems connected with the weldability of Inconel 713C. The studies included analyses of the macro- and microstructure of the clad welds, the base materials, and the heat-affected zones. The results of the structural analyses of the clad welds indicate that Inconel 713C should be classified as a low-weldability material. In the clad welds made by laser, cracks were identified mainly in the heat-affected zone and at the melted zone interface, crystals were formed on partially-melted grains. Cracks of this type were not identified in the clad welds made using the plasma-arc method. It has been concluded that due to the possibility of manual cladding and the absence of welding imperfections, the technology having the greatest potential for application is plasma-arc cladding.

Cycle Optimization Potential of Composite Cycle and Turbocompound Aero-Engines

2020 ◽  
Vol 142 (6) ◽  
Author(s):  
Felix Klein ◽  
Stephan Staudacher
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Exchange Process ◽  
Aircraft Engine ◽  
Time Frame ◽  
Fuel Burn ◽  
Component Efficiency ◽  
Low Efficiency ◽  
Aero Engines

Abstract Fair comparison of future aircraft engine concepts requires the assumption of similar technological risk and a transparent book keeping of losses. A 1000 km and a 7000 km flight mission of a single-aisle airplane similar to the Aribus A321neo LR have been used to compare composite cycle engines, turbocompound engines and advanced gas turbines as potential options for an entry-into-service time frame of 2050+. A 2035 technology gas turbine serves as reference. The cycle optimization has been carried out with a peak pressure ratio of 250 and a maximum cycle temperature of 2200 K at cruise as boundary conditions. With the associated heat loss and the low efficiency of the gas exchange process limiting piston component efficiency, the cycle optimization filtered out composite cycle concepts. Taking mission fuel burn (MFB) as the most relevant criterion, the highest MFB reduction of 13.7% compared to the 2035 reference gas turbine is demonstrated for an air-cooled turbocompound concept with additional combustion chamber. An intercooled, hectopressure gas turbine with pressure gain combustion achieves 20.6% reduction in MFB relative to the 2035 reference gas turbine.

Sign in / Sign up

Export Citation Format

Share Document