Numerical Study and Design of an Interstage Turbine Burner

Siow, Yeow; Yang, Shu

doi:10.2514/6.2002-3727

Cited by 7 publications

(7 citation statements)

References 4 publications

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“…The name Interstage Turbine Burner was coined by Siow and Yang [15] to describe the miniature combustor in the transition duct between a high-pressure turbine section and a low-pressure turbine section. It should be noted that all the other cycle studies presented so far assumed a single-spool engine, where all the compressor stages were driven through a single shaft by all the turbine stages.…”

Section: Cycle Studies On Engines With Interstage Turbine Burnermentioning

confidence: 99%

“…To apply this design to an aircraft engine, however, the secondary combustor must be much smaller than any current combustor. As a result, Siow and Yang [15] proposed the Interstage Turbine Burner cycle based on the miniature combustor they studied, even though the cycle had been investigated to a certain degree in other researches.…”

Section: Figure 23 a Schematic Comparison Between An Engine With Andmentioning

confidence: 99%

“…It was demonstrated that combustion can be enhanced by the introducing swirl to the flow or by the presence of centrifugal force on the flow. While earlier works on this effect were more geared toward designing a more efficient combustor, several later researches, including the study by Siow and Yang [15] discussed earlier, proposed to design compact combustors using the results from earlier works. These compact or miniature combustors are quite different from the concept of a turburner, but the phenomenon of enhanced combustion in a swirling flow is certainly worth discussing as the centrifugal force and swirling in a turbine flow is very high.…”

Section: Numerical and Experimental Studies On A Miniaturized Combustormentioning

confidence: 99%

“…where π d is the overall total pressure ratio across the entire inlet system, π dmax is the total pressure ratio caused by wall friction effects alone, and η R is the ram recovery factor. For the value of π dmax , it is determined empirically and chosen based on the recommendations outlined in the work of Mattingly, Heiser, and Daley 15 . The ram recovery factor is based on military specification MIL-E-5008B [25], as given by ( )…”

Section: Module For Freestream and Inletmentioning

confidence: 99%

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A Performance Study of a Super-Cruise Engine with Isothermal Combustion Inside the Turbine

Chiu¹

2005

41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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Current thinking on the best propulsion system for a next-generation supersonic cruising (Mach 2 to Mach 4) aircraft is a mixed-flow turbofan engine with afterburner. This study investigates the performance increase of a turbofan engine through the use of isothermal combustion inside the high-pressure turbine (High-Pressure Turburner, HPTB) as an alternative form of thrust augmentation. A cycle analysis computer program is developed for accurate prediction of the engine performance and a supersonic transport cruising at Mach 2 at 60,000 ft is used to demonstrate the merit of using a turburner. When assuming no increase in turbine cooling flow is needed, the engine with HPTB could provide either 7.7% increase in cruise range or a 41% reduction in engine mass flow when compared to a traditional turbofan engine providing the sane thrust. If the required cooling flow in the turbine is almost doubled, the new engine with HPTB could still provide a 4.6% increase in range or 33% reduction in engine mass flow. In fact, the results also show that the degradation of engine performance because of increased cooling flow in a turburner is less than half of the degradation of engine performance because of increased cooling flow in a regular turbine. Therefore, a turbofan engine with HPTB will still easily out-perform a traditional turbofan when even more cooling than currently assumed is introduced. Closer examination of the simulation results in off-design regimes also shows that the new engine not only satisfies the thrust and efficiency requirement at the design cruise point, but also provides enough thrust and comparable or better efficiency in all other flight regimes such as transonic acceleration and takeoff. Another finding is that the off-design bypass ratio of the new engine increases slower than a regular turbofan as the aircraft flies higher and faster. This behavior enables the new engine to maintain higher thrust over a larger flight envelope, crucial in developing faster air-breathing aircraft for the future. As a result, an engine with HPTB provides significant benefit both at the design point and in the off-design regimes, allowing smaller and more efficient engines for supersonic aircraft to be realized.

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Section: Cycle Studies On Engines With Interstage Turbine Burnermentioning

confidence: 99%

Section: Figure 23 a Schematic Comparison Between An Engine With Andmentioning

confidence: 99%

Section: Numerical and Experimental Studies On A Miniaturized Combustormentioning

confidence: 99%

Section: Module For Freestream and Inletmentioning

confidence: 99%

See 2 more Smart Citations

A Performance Study of a Super-Cruise Engine with Isothermal Combustion Inside the Turbine

Chiu¹

2005

41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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show abstract

“…A deciding factor is the time required for mixing and burning the fuel. An alternate strategy being investigated 11,12 is to incorporate an inter-stage burner in the transition duct between the high-and low-pressure turbines. These and the current investigation suggest that further research is warranted to determine an optimal implementation strategy for various applications.…”

Section: Overall Comparisons Of the Three Cooling Conceptsmentioning

confidence: 99%

Gas-Turbine Performance Improvements Through the Use of Multiple Turbine Interstage Burners

Chen

Hoffman

Davis

2004

Journal of Propulsion and Power

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Performance evaluation of a novel re-cooled mixed-flow turbofan cycle for aviation power application

Sun

Wen

et al. 2020

J Therm Anal Calorim

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A novel re-cooled mixed-flow turbofan cycle (RMTC) is proposed to meet the demand of high thrust-weight ratio for military aero-engine. Through adding a re-cooler during the compression process, the specific thrust can be improved by increasing the outer fan flow temperature and also the turbine cooling air consumption can be reduced by decreasing the inner fan flow temperature. This work conducts the modeling of ideal and actual RMTCs by using a verified original program. The effects of re-cooler location, heat transfer effectiveness, and bypass ratio on the specific thrust and fuel consumption rate are evaluated, and the RMTC performances in the entire flight envelope are discussed to estimate its applicable flight conditions. Based on the parametric analysis, locating the re-cooler after the compressor and increasing the heat transfer effectiveness are beneficial for the increase in specific thrust, while the effect of bypass ratio is related to the heat transfer capacity of recooler. The specific thrust improvement of RMTC is more significant at high flight Mach number, and hence, larger aircraft payload can be obtained with the same fuel mass. The assessment results show that the proposed RMTC system exhibits potential application for high-speed military aircraft.

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Numerical Study and Design of an Interstage Turbine Burner

Cited by 7 publications

References 4 publications

A Performance Study of a Super-Cruise Engine with Isothermal Combustion Inside the Turbine

A Performance Study of a Super-Cruise Engine with Isothermal Combustion Inside the Turbine

Gas-Turbine Performance Improvements Through the Use of Multiple Turbine Interstage Burners

Performance evaluation of a novel re-cooled mixed-flow turbofan cycle for aviation power application

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