A challenge in establishing rotor performance map for sizing tool during design cycle is the rotor performance uncertainty for full vehicle. Sometimes, simplified tests at different setup/scale are conducted to guide performance map, but this introduces another uncertainty due to configuration difference from full vehicle. To aid insights, validated computational fluid dynamics simulations (using CREATE-AV™ Helios) were carried out to examine hovering rotor performance prediction variations at different design stages, or different modeling/testing setup with identical blade design. Quantitative rotor figure of merit differences has been demonstrated along with descriptions of underlying physical reasons. The examined model setup includes isolated rigid blades with and without flapping, elastic blades, model-scale blades, whirl-tower conditions, blades installed on fuselage, and full-vehicle including tail rotor. Both fully turbulent flow and laminar-turbulence transition flow assumptions were simulated. Rigid blades showed a negative impact on performance due to the lack of nose-down elastic twist. Model scale suffered from lower Reynolds number effects but took advantage of delayed laminar-turbulence transition. Whirl-tower blockage and altered vortex trajectories reduced peak figure of merit but delayed stall. Fuselage installation increased performance with partial in-ground effect. However, the tail rotor disrupted the main rotor vortex system and caused a substantial figure of merit drop. The figure of merit variation summary from current study can provide qualitative trends and rough estimate of the rotor performance change along different analysis or test condition during new design process. Well validated computational fluid dynamics simulations can be used as a risk-reduction approach by comparing with simplified-model results (used during fast design cycle) with the expected full-vehicle model results.

Impact of laminar-turbulent flow transition to hovering rotor performance is assessed using both model- and full-scale rotors. Helios is used for all simulations using OVERFLOW as a near-body solver. The Langtry-Mentor transition model is validated against Hover Validation and Acoustic Baseline (HVAB) rotor test data using industry best practice modeling approach. The same modeling approach is validated with model-scale production design rotor to assess its applicability to different blade design including different airfoils, twist, and planform shape. The validated modeling approach is then used for full-scale rotor to assess impact of transition for high Reynolds-number full-scale rotor performance. The full-scale rotor performance is also compared with whirl-tower test data. For all three rotors, rotor figure of merit is compared for range of thrust against test data. Transition locations and blade loading distributions are also compared. Sensitivity studies were also performed including facility impact, control system stiffness effect, and wind impact. The current transition modeling approach showed very good correlation with both model-scale test data in integrated figure of merit. However, differences with test data were observed in detailed blade loading and surface suction pressure at outboard region where blade-vortex interaction occurs. The full-scale rotor simulation also showed reasonably good correlation in performance but with simulation results being slightly optimistic in both fully turbulent and transition model, with transition model being little more optimistic. The laminar-turbulent flow transition trend from model- to full-scale seems reasonable, although, with more uncertainties in the full-scale test results, it was hard to determine quantitative transition model impact validation. Findings from current study is used to guide future hover performance simulations, and to understand model- and full-scale rotor performance difference.

The Langtry-Mentor transition model with k-omega Shear Stress Transport (SST) model has been assessed for multiple rotors. In this assessment, an adverse impact of the sustaining term has been identified and a modification has been introduced. It was found that the original sustaining term behaved poorly for the separated flows investigated herein, making separation worse and leading to a premature stall in hover. The modified sustaining term uses F1-mu(t) blending to deactivate the sustaining term inside boundary layer. The Langtry-Mentor transition model with the modified sustaining term has been validated for multiple rotors with various scales and characteristics, including the XV-15 rotor, the Hover Validation and Acoustic Baseline (HVAB) rotor, and Sikorsky's model- and full- scale rotors. Helios, with OVERFLOW as near-body solver, was used for these calculations. The validation showed very good correlation with measured hover figure of merit data for all rotors from low thrust to high thrust conditions, using a consistent transition model and sustaining-term settings. It was found that the modified sustaining term removed the adverse impact of the original sustaining term.

Full vehicle CREATE-AV™ Helios model has been developed for SB>1 Defiant®, a Joint-Multi-Role Technology Demonstrator (JMRTD) designed by the Sikorsky-Boeing team utilizing compound design with coaxial rotor and propulsor. The full vehicle model includes coupling with RCAS for elastic blade deformation as well as full vehicle trim in steady level flight condition. The purpose of current study is to assess performance prediction capability of the developed Helios modeling approach for such a complex non-traditional design by correlating with flight test data. To minimize uncertainty in flight test data reduction, correlations were made with the data that were directly measured or requiring minimal derivation. The Helios model showed generally very good correlation in power, component forces, rotor and propulsor efficiencies for wide range of flight test conditions. The model also showed very good correlation in performance sensitivity to trim state and rotor RPM, which demonstrates the modeling approach can be used to find optimal flight trim condition. Significance of geometric details and its aerodynamic interference such as shaft modeling has been demonstrated. Impact of measurement uncertainty to data correlation was also demonstrated.

Interactional aerodynamic interactions between various rotorcraft components can make a large contribution to steady and unsteady loads, performance, and vibration. Wind tunnel results from a powered model test have been analyzed to identify trends in the unsteady aerodynamic pressures on the horizontal stabilizer. Flow velocity measurements were also made behind the fuselage, rotor hub, and blades. Velocity components in all three directions were separated into time-averaged, periodic, and broadband components to identify factors contributing to unsteady tail loads and provide validation data for analysis. Computational Fluid Dynamics (CFD) has been applied to four configurations of the wind tunnel model. The calculated steady rotor and fuselage forces and the unsteady tail pressures have been compared to experiment. CFD has also been applied to a flight test configuration and the results compared to measured stabilizer accelerations. When all relevant components are included, the CFD analysis captures many key features, but there remains room for improvement in resolving the quantitative details.

CREATE-AV Helios/Overflow hover simulations were performed for the S-76® model-scale rotor out-of-ground effect. The two blade planforms with blade tip variations that are investigated in this study include the swept-tapered and swept-tapered-anhedral tips. Performance calculations for the two S-76 blade planforms were compared with the publicly available experimental data set. Sensitivity to coning angle were investigated, and sensitivity to light wind conditions. In addition, full-scale rotor hover simulations were performed for out-of-ground effect conditions, and with the tower/ground effects included to compare directly with full-scale whirl-test results. STAR-CCM+ simulations were also performed and compared with full-scale test results and HELIOS, for both in-ground-effect and out-of-ground effect. Correlation with full-scale results is very good, when the integration effects are included. Model-scale results tend to be low in Figure of Merit versus test using a baseline mesh for the AIAA hover workshop. Full-scale rotor calculations used a finer mesh and provided good correlation with test.

The CH-53K® King Stallion™ is the most advanced heavy lift helicopter developed by Sikorsky, a Lockheed Martin Company, to address the requirements of the United States Marine Corps. The aircraft was designed to support missions with a maximum design gross weight of 88000 lbs and can carry external loads up to 36000 lb. Full-aircraft HELIOS calculations were performed for a CH-53K King Stallion including main rotor, tail-rotor, fuselage and detailed hub. Calculations were performed for various configurations, with main-rotor removed, tail-rotor removed, and isolated main and tail rotors to evaluate the multiple interactional effects. Comparisons of the total power show excellent agreement with the flight-test measurements.

Current paper summarizes a correlation study of two flow solvers (CREATE-AV™ Helios and STAR-CCM+), routinely used at Sikorsky, with the spinning coaxial hub drag and flow field measurements conducted by Penn State University at the 12' diameter water tunnel. The Helios modeling approach was aiming for a high fidelity accurate simulation, whereas the STAR-CCM+ modeling approach was aiming for a fast turn-around time with reasonable solution accuracy with a relatively coarse mesh and simplification. The two solvers generally agreed well with the test data within reasonable accuracy and captured the drag trend between two shaft fairing configurations. Impact of turbulence model selection (Spalart-Allmaras Detached Eddy Simulation and Spalart-Allmaras Reynolds-Averaged-Navier-Stokes model) has been demonstrated. The RANS model generally delayed separation and resulted in lower drag. The STAR-CCM+ runs simulated both air and water at matching Reynolds number and showed good agreement between the drag results for the two mediums. Also, the importance of accurate representation of geometric details including gaps, shafts, and holes is highlighted.

The aeroelastic design of a high speed highly efficient rotor, from conceptual layout to detailed design, is presented. The overall strategy is described along with details and a demonstration case on the conceptual design optimization strategy. In Phase I, rapid design iterations were performed using lower order quick turnaround tools to establish basic design aspects such as rotor head sizing and stiffnesses. In Phase II, higher order tools are implemented along with a more realistic structural representation of the head and blades. The outcome is a robust preliminary design to be analyzed with CFD-CSD and tuned further in detailed design without requiring major rework. Finally, dynamic tailoring performed on the blade during final design is presented. This was done to maintain or improve frequency placements and vibration levels as relatively small, surgical changes were made to the blade detailed design (ply thicknesses, orientations, blade weights, etc.) to satisfy strength, life, and manufacturability requirements. This work highlights the importance of upfront aeromechanics optimization to establish a robust early conceptual design, and high-fidelity analysis through final design to maintain desired characteristics.

Developmental flight testing on the United States Marine Corps (USMC) CH-53K King Stallion heavy lift program identified significant engine integration technical challenges including Exhaust Gas Reingestion (EGR), engine compartment cooling and engine exhaust impingement on the fuselage. These challenges limited the mission envelope and threatened to delay the CH-53K King Stallion heavy lift helicopter program. Immediate resolution of these issues was required for the USMC to proceed with the planned deployment of CH-53K. The CH-53K all digital design, advanced computer modeling techniques, rapid prototyping and flight test data reduction methodologies were brought together by the U.S. Navy and Industry team to solve these problems and progress the program. This paper will cover multiple facets of the CH-53K flight testing conducted to investigate and characterize these complex, interrelated issues and then used to demonstrate the robustness of the selected design after parts were delivered to the flight vehicle. This paper will discuss the up-instrumentation effort required during flight test. It will also cover the subsequent data reduction, analysis, and visualization techniques implemented to understand and illustrate flight test results and accurately characterize Exhaust Gas Reingestion.

When a helicopter hovers close to the ground, the flow field around rotor and fuselage is highly complex with the rotor down-wash interacting with the ground and fuselage. The reflection effect of the ground plane can cause an upwash through the rotor inboard area causing a highly unsteady recirculation zone. One of critical problems that can be caused by this interactional effect is engine exhaust gas re-ingestion, which potentially can cause engine stall. These interactions are further complicated by the presence of light winds. The current paper summarizes physical mechanisms of exhaust gas re-ingestion of a heavy lift helicopter analyzed by numerical simulations and flight test data. Tools in different categories of computational fluid dynamics are validated against measured data. The contribution from each exhaust to EGR and its mechanism were identified at different wind directions.

Developmental flight testing on the United States Marine Corps (USMC) CH-53K King Stallion heavy lift program identified significant engine integration technical challenges including Exhaust Gas Reingestion (EGR), engine compartment cooling and engine exhaust impingement on the fuselage. These challenges limited the mission envelope and threatened to delay the CH-53K King Stallion heavy lift helicopter program. Immediate resolution of these issues was required for the USMC to proceed with the planned deployment of CH-53K. The CH-53K all digital design, advanced computer modeling techniques, rapid prototyping and flight test data reduction methodologies were brought together by the U.S. Navy and Industry team to solve these problems and progress the program. This paper will provide an overview of the engine integration challenges, model based design approach employed by the team towards most extensive CFD modeling effort exercised to-date across the rotorcraft industry, largest HPC utilization to-date at Sikorsky and how the high fidelity modeling effort, rapid prototyping and advanced flight test data reduction techniques helped in the mitigation of EGR and engine integration technical challenges on CH-53K

Developmental flight testing on the United States Marine Corps (USMC) CH-53K King Stallion heavy lift program identified significant engine integration technical challenges including exhaust gas re-ingestion, engine compartment cooling and engine exhaust impingement on the fuselage. These challenges limited the mission envelope and threatened to delay the CH-53K King Stallion heavy lift helicopter program. Immediate resolution of these issues was required for the USMC to proceed with the planned deployment of CH-53K. The CH-53K all digital design, advanced computer modeling techniques, rapid prototyping and flight test data reduction methodologies were brought together by the U.S. Navy and Industry team to solve these problems and progress the program. One of the key considerations when modifying an engine exhaust system, particularly for helicopters which operate in a wide range of flight conditions (forward, side, and rear flight, for example), is the impact to the performance of the exhaust and engine bay cooling. This paper focuses on the engine bay cooling modeling aspects of the design investigation as well as CFD validation and parallel flight test support to mitigate bay cooling limitations.

Developmental flight testing on the United States Marine Corps (USMC) CH-53K King Stallion heavy lift program identified significant engine integration technical challenges including Exhaust Gas Reingestion (EGR), engine compartment cooling and engine exhaust impingement on the fuselage. These challenges limited the mission envelope and threatened to delay the CH-53K King Stallion heavy lift helicopter program. Immediate resolution of these issues was required for the USMC to proceed with the planned deployment of CH-53K. The CH-53K all digital design, advanced computer modeling techniques, rapid prototyping and flight test data reduction methodologies were brought together by the U.S. Navy and Industry team to solve these problems and progress the program. This paper details the final design configuration of the CH-53K – the D124 configuration. This paper will focus on the final exhaust modifications, D124 and D83, proposed through extensive CFD analysis performed by the integrated Sikorsky, NAVAIR, Army, and General Electric (GE) team. The paper will detail the successful demonstration of the design solution, “D124”, through extensive flight testing. In addition, this paper will include design optimizations of the #2 exhaust duct that continued after the D124 flight test. T he analytical methodology utilized to assess robustness of the design modifications across a large mission envelope will be detailed. The resultant process established a fly-analyze-design-fly approach and enabled a rapid turnaround time from problem identification to solution implementation with high confidence and low risk. CFD and flight test data were complementary throughout the effort. Used together they enabled extensive validation of the CFD, an increased comprehension of the empirical results and the ability to virtually expand the design space. The paper will describe how the team successfully demonstrated their capability to replace the traditional long-lead, serial and schedule intensive ‘fly-fix-fly-fix’ design approach to resolve complex, flight test discoveries with a high confidence approach guided by high fidelity modeling in a compressed schedule. This paper takes a detailed look at the final CFD recommendations that resulted from the trade study and contrasts the downselected exhaust configuration, D124, impact relative to the CH-53K baseline exhaust configuration. Details of the extensive testing performed to assess the ability of the D124 design modifications to address EGR, engine bay cooling and impingement will be described. Flight test results in worst case maneuvers indicated a greater than 50% reduction in inlet distortion and engine inlet temperature rise versus the baseline configuration. Flight test results also indicated that the EGR reduction also addressed the prior deficiency of EGR sensitivity to aircraft orientation with respect to wind direction. Reduced EGR in the modified design was demonstrated to be invariant to aircraft orientation with respect to wind direction, significantly enhancing the robustness of the CH-53K aircraft. Reduction of EGR addressed the root-cause for CH-53K engine stalls. In addition to EGR, the design concept mitigated engine compartment cooling and engine exhaust impingement challenges across a range of flight profiles in the mission envelope of CH-53K. The pre-test CFD predictions for critical engine operability and integration metrics will be compared to the D124 flight test data. This paper will also provide a final discussion of the cadence of the EGR investigation, including the large strides taken to reduce CFD lead time and the implication and benefits toward air vehicle design. The paper will describe lessons learned along the way that have application to the entire aerodynamic industry and its use of CFD.

Developmental flight testing on the United States Marine Corps (USMC) CH-53K King Stallion heavy lift program identified significant engine integration technical challenges including Exhaust Gas Reingestion (EGR), engine compartment cooling and engine exhaust impingement on the aircraft’s fuselage. These challenges limited the mission envelope and threatened to delay the CH-53K King Stallion heavy lift helicopter program. Immediate resolution of these issues was required for the USMC to proceed with the planned deployment of CH-53K. The CH-53K all digital design, advanced computer modeling techniques, rapid prototyping and flight test data reduction methodologies were brought together by the U. S. Navy and industry team to solve these problems and progress the program. This paper discusses the work conducted in parallel and immediately after the Computational Fluid Dynamics (CFD) modeling validation effort; CFD was deployed in a massive way in order to examine the air vehicle impacts of different EGR solution proposals at both critical EGR flight conditions as well as across the CH-53K flight envelope. CFD data extractions, focused on EGR attributes and engine integration risks, were then used in a multi-disciplinary systems engineering trade study that weighted pertinent air vehicle impacts, including, e. g. , aircraft performance, weight, engine life, cost of implantation and human factors. The system engineering score was used to downselect high confidence aircraft configurations for flight test assessment.

The S-97 RAIDER™ is a next-generation light tactical helicopter which uses a coaxial dual main rotor and aft-mounted propeller for auxiliary propulsion. This advanced configuration presents challenges for aerodynamic modeling, particularly with regard to the interactions between various components and their dependency on flight condition and trim state across a wide envelope. This paper describes continued research into the aerodynamics of the coupled aircraft components. Computational simulations of multiple combinations of components have been conducted for a series of pitch sweeps and flight conditions. These computational fluid dynamics (CFD) results are compared to available scale-model wind-tunnel test results. As is typical with development programs, changes were made between the original small-scale testing and analysis, and the current flight-test aircraft. In addition, the available computational methods and processing capabilities have improved significantly since the program started in 2010.

This study is focused on the improvement of UH-60A blade structural loads correlation with flight and wind-tunnel test data. The blade airloads prediction has been proven to be reasonably good from past studies. However, the blade structural loads, especially the edgewise bending moment, were poorly predicted, and these are the subject of the current study. Several variations of modeling effort have been examined. This includes the drive-train model, refined grid, lag-damper model variation, hub impedance with test stand, tunnel wall, and dis-similar blades. The drive-train and refined grid showed noticeable improvement in edgewise bending moment, but not enough to close the gap between measured data and prediction. A lag damper study indicated that the current nonlinear damper seems to be adequate, and variations to the lag damper had a limited impact to inboard region only. The influence of dis-similar blades provides further insight into the sensitivity of the edgewise bending moment. The largest difference was observed to be due to the presence of the fuselage-like wind tunnel mount, indicating that a rotor analysis should include the fuselage.