Rotorcraft Development Research Articles

Exploration of stacked rotor designs for aerodynamics in hover

The development of rotorcraft with multiple propulsors has increased interest in co-rotating coaxial rotors, also known as stacked rotors, which can generate high thrust while maintaining a compact aircraft design. By optimizing the pitch angle and the blade arrangement, stacked rotors can reduce aeroacoustic noise and vibration while improving aerodynamic performance. This study leveraged a high-fidelity numerical solver to analyze the aerodynamic performance and underlying physics of various stacked rotor configurations in hovering conditions. With three rotor design variables — index angle, stacked distance, and pitch angle difference — together with high-resolution simulations, two key physical phenomena were identified: the inflow effect and the blade-vortex interaction effect. Specifically, the blade-vortex interaction at the lower blade was found to play a critical role in improving hover performance by generating upwash, which increased the local power loading of the lower blade. Also, the unsteady simulations revealed that stacked rotors produced significantly lower peak-to-peak thrust values compared to counter-rotating coaxial rotors, indicating potential reductions in aeroacoustic noise level. Finally, a design optimization based on deep neural networks was performed to successfully extract design rules to take full advantage of inflow and blade-vortex interaction effects, which can be used as a guideline for the future design study of stacked rotors.

Mechanical Analysis Process of a Coaxial Counter Rotor for Applications in Unmanned Ultra-Light Units

Abstract As the rotor configuration has the most impact on helicopter properties, the process of determination the assumptions for rotor design is a very important factor in the early stage of rotorcraft development. The following paper presents a mechanical analysis process used at the Institute of Aviation to quickly develop a coaxial rotor prototype applicable in ultra-light unmanned helicopter which has the potential for further improvement of its flight parameters. The article describes the rotor analysis process due to its feasibility based on commercially available solutions, the process of formulating assumptions for the entire structure, MES analysis of the rotor parts all leading to creation of the rotor prototype.

Predicting the Influence of Blade Shape on Hover Performance with Comprehensive Analyses

The ability to predict and understand the influence of rotor tip shape on hovering rotor performance is critical to rotorcraft development. Comprehensive rotorcraft analysis methods are an efficient way to reduce the risk and cost associated with developing new rotor designs. In the modern design environment, these methods are ideal for efficiently traversing the design space in order to identify designs of interest suitable for more detailed analysis, testing, and optimization. In this paper, the ability of three distinctly different comprehensive rotorcraft methods (EHPIC, CHARM, and VTM) to predict the hover performance of a rotor with a variety of tip shapes is evaluated. The assessment indicates that good predictive fidelity is achieved for the cases considered, though dependence on airfoil data remains a limiting factor in the generality of the analysis results.

Numerical calculation of energy and exergy flows of a turboshaft engine for power generation and helicopter applications

Fuel efficiency of aircraft and helicopter becomes greater concern in recent years caused by rising fuel costs and as well as environmental impact of aviation emissions. Modern helicopters, however, highly complex systems with especially turboshaft engines that produce energy and power. So it is important to gain deeper understanding energy and exergy use throughout turboshaft engine and its components. Concurrently, in this study, energy and exergy-based computational approach applied to a turboshaft engine and its components. Exergy efficiency of the axial, centrifugal compressors and power turbine is found to be between 83.8% and 88.6%, while for the combustor, the corresponding value is to be 80.60%. For the components, the greatest exergy efficiency is calculated to be 91.4% at the gas generator turbine unit. As a result of the study, the exergetic efficiency of the turboshaft has been calculated as 27.5% with 1500 kW product exergy. It is expected that numerical formulation based on energy and exergy is beneficial for assessing turboshaft performance for future rotorcraft development.

항공기용 연료탱크 Phase I 충돌충격시험 충격하중 분석

Crash impact test is conducted to verify the crashworthiness of fuel tank. Success of the crash impact test means the improvement of survivability of crews by preventing post-crash fire. But, there is a big risk of failure due to huge external load in the crash impact test. The failure of crash impact test can result in serious delay of a entire rotorcraft development because of the design complement and re-production of the test specimens requiring a long-term preparation. Thus, the numerical simulations of the crash impact test has been required at the early design stage to minimize the possibility of trial-and-error in the real test. Present study conducts on the numerical simulation of phase I crash impact test using SPH supported by crash simulation software, LS-DYNA. Test condition of MIL-DTL-27422 is reflected on analysis and material data is acquired by specimen test of fuel cell material. As a result, the crash load on the skin material, overlap area and metal fitting is estimated to confirm the possibility of acquisition of the design load for the determination of the overlap area and adhesive strength.

회전익항공기 연료셀 충돌충격시험 Full-Scale 수치모사

항공기 연료셀은 추락 상황에서 승무원의 생존성과 직결되는 중요 구성품으로 회전익 항공기에 적용되고 있는 내충격성 연료셀은 추락시 승무원의 생존성 향상에 큰 역할을 하고 있다. 미육군은 항공기가 처할수 있는 다양한 상황에서 연료셀이 제 기능을 발휘할 수 있도록 1960년대 초부터 MIL-DTL-27422 이라는 연료셀 개발규격을 제정하여 현재까지 적용해 오고 있다. 해당 개발규격에 규정된 시험 중에서 충돌충격시험은 연료셀의 내충격 성능을 검증하는 시험으로써, 해당 시험을 통과하는 연료셀은 생존가능 충돌환경에서 화재가 발생하지 않아 승무원의 생존성이 대폭 향상될 수 있음을 의미한다. 그러나 충돌충격시험은 작용하는 하중 수준이 너무 높기 때문에 실패 위험성이 가장 큰 시험이기도 하다. 연료셀이 해당 시험을 통과하지 못하는 경우에는 재시험을 위한 비용과 준비기간이 상당히 소요되어 항공기 개발일정에 심각한 지장을 초래할 가능성도 높다. 따라서, 연료셀 설계 초기부터 내충격성능 만족여부에 대한 예측을 위해 충돌충격시험의 수치해석을 통한 실물시험에서의 실패 가능성을 최소화해야 한다는 필요성이 제기되어 왔다. 본 연구에서는 충돌모사 프로그램인 LS-DYNA에서 지원하는 유체-구조 연성해석 방법인 SPH 방법을 사용하여 연료셀 충돌충격시험 수치 모사를 수행하였다. 수치해석 조건으로 MIL-DTL-27422에서 요구하는 시험조건을 고려하였고, 실물 연료셀의 시편시험을 통해 확보한 물성데이타를 해석에 반영하였다. 그 결과로 연료셀 자체의 응력수준을 평가하고 취약부위에 대한 고찰을 수행하였다. Crashworthy fuel cells have a great influence on improving the survivability of crews. Since 1960's, the US army has developed a detailed military specification, MIL-DTL-27422, defining the performance requirements for rotorcraft fuel cells. In the qualification tests required by MIL-DTL-27422, the crash impact test should be conducted to verify the crashworthiness of fuel cell. Success of the crash impact test means the improvement of survivability of crews by preventing post-crash fire. But, there is a big risk of failure due to huge external load in the crash impact test. Because the crash impact test itself takes a long-term preparation efforts together with costly fuel cell specimens, the failure of crash impact test can result in serious delay of a entire rotorcraft development. Thus, the numerical simulations of the crash impact test has been required at the early design stage to minimize the possibility of trial-and-error with full-scale fuel cells. Present study performs the numerical simulation using SPH(smoothed particle hydro-dynamic) method supported by a crash simulation software, LS-DYNA. Test condition of MIL-DTL-27422 is reflected on analysis and material data is acquired by specimen test of fuel cell material. As a result, the resulting equivalent stresses of fuel cell itself are calculated and vulnerable areas are also evaluated.

Fly-By-Wire 헬리콥터 비행제어법칙 자동생성코드 개발 및 검증 프로세스

전자식 비행제어시스템(Fly-By-Wire Flight Control System)에 적용되는 헬리콥터의 제어법칙 설계 및 해석은 모델기반 설계기법(Model Based Design)을 사용하여 개발되고 있다. 모델기반 설계기법은 회전익뿐만 아니라 고정익, 무인항공기, 자동차 산업에서도 널리 사용되고 있다. 모델기반 설계기법은 핵심기술인 시뮬레이션과 자동생성코드를 이용하여 개발기간 단축 및 개발리스크 감소라는 장점을 제공한다. 본 논문에서는 FBW 헬리콥터 비행제어법칙의 자동생성코드에 대해서 체계적인 기능검증을 위한 프로세스를 개발하였다. 그리고 지식경제부 과제로 진행하고 있는 FBW 헬리콥터 개발에 적용하여 비행제어법칙 자동생성코드의 기능 검증을 수행 결과, 요구 조건을 충족하는 것으로 확인할 수 있었다. The control law design and analysis environment of the FBW helicopter system have been developed using model base design method. The model based design is generally used in a aircraft, unmanned aerial system and automobile as well as rotorcraft development. The model based design provides many advantages such as development risk and schedule reduction using simulation and autocode generation. This paper describes a development of process for verification and validation of auto generated code for FBW helicopter flight control law. And this process is applied for Fly-By-Wire Helicopter Development Project. The results of functional test for auto generated code meet several specific requirements.

Some Aspects of Future Helicopter and other Rotorcraft Development

VERTICAL take‐off aircraft in the form of jump‐autogyros and helicopters have been under sporadic development since the first world war during which a co‐axial design attributed to von Karman was used tethered for observation purposes. Only limited success was achieved between the wars due to two main factors:

Frequency and Vibration Problems of Rotors

I proceeding from fixed wing aircraft to rotating wing aircraft, there are not only new and complicated aerodynamic and stability problems but also new vibration problems to be met and solved. These vibration problems have undoubtedly been a major item in retarding the development of rotorcraft. Theoretical studies of rotorcraft vibration and frequency problems, which were initiated about two years ago by the Kellett Autogiro Corporation, have culminated in this paper. This study has been confined to the frequency problems which are peculiar to hinged rotor blades. It is felt that rigid-blade rotor construction may be analyzed by substantially the same methods as are available for analyzing propellers and, therefore, no consideration has been given to this specialized type in this paper. When a rotorcraft is operating in horizontal flight, the airflow over the blades on the advancing side of the rotor is considerably greater than the airflow over the retreating side. This unsymmetrical flow of air over the blades, in combination with the non-uniform flow of air through the rotor disc, causes the blades to rise and fall during each revolution of the rotor. In addition, the blades flex in bending in the plane of flapping and oscillate about the vertical hinges with a frequency of one cycle for each revolution of the rotor.