Browse Topic: Thrust
Gaganyaan is an ambitious and recover safety mission for the Indian space program to launch humans into space. The success of the mission depends on the development of required technology and systems. A test vehicle is developed for the technological demonstration for all envisioned abort flight scenarios of Gaganyaan mission. A new configuration of launch vehicle with single liquid stage is planned for multiple flights. Coupled Loads analysis of launch vehicle system is a standard practice to estimate response and loads for the design of structures and generating sine vibration test levels. Usually a vehicle rests on the launch pad through base shroud with horizontal support and no vertical restraint. Upon ignition of the engine, thrust builds up and upon overcoming gravity the vehicle takes off. In the current analysis the launch vehicle is held in position using a holding / retracting mechanism and at a predefined time the vehicle is released. The boundary condition required a novel
This paper explores the effect of addition of a horizontal tail on the longitudinal stability and performance of a Biplane Tailsitter Unmanned Aerial Vehicle (UAV). Biplane tailsitters a type of hybrid UAVs, often exhibits poor longitudinal stability during forward flight, necessitating continuous active control through application of differential motor thrust to maintain attitude. To address this challenge, this work proposes the integration of a horizontal tail on a quadrotor biplane tailsitter UAV, aiming to improve pitch stability and control authority during critical flight phases. Experimental flight data was utilized to determine the appropriate sizing of the elevator. A detailed flight dynamics model validated the effectiveness of the elevator control. The design was validated through outdoor flight testing, comparing the performance of tail-less and tail-attached configurations. The results demonstrate that the modified design results in a reduction control power requirement
This paper presents experimental research aimed at developing novel low lubrication methods for rotorcraft and jet engines, focusing on sustaining minimal lubrication to prevent catastrophic bearing failure during loss of lubrication (LoL) events or to increase fuel consumption performance on once-through, fuel-oil bearing lubrication engines. Utilizing two high-speed bearing test rigs simulating low and high thrust class engine conditions, the study establishes lower bounds for oil flow rates necessary to maintain thermal stability and prevent thermal runaway in hybrid ball bearings. These findings inform the design of the Zulu Pod (ZPod), a passively driven, self-contained oil delivery system that uses engine compressor bleed air to precisely meter lubricant flow. Engine test stand results demonstrate that replacing traditional fuel-oil lubrication with the ZPod system reduces thrust specific fuel consumption (TSFC) by an average of 7%, with up to 11% savings, without compromising
A wind tunnel investigation to characterise the aerodynamic performance and aeroelastic response of a tiltrotor blade set operating in propeller mode is presented. A custom blade set was instrumented with fully bridged axial strain gauges to monitor the flap bending and torsional strain at several radial locations. Propeller thrust and torque measurements were acquired using a custom six component Rotating Shaft Balance. Measurements of blade tip deflection were obtained via stereoscopic Digital Image Correlation. Testing was performed at a range of rotational frequencies, blade pitch angles and advance ratios to assess the blade aerodynamic performance and aeroelastic response in both attached and stalled operating conditions. Strain measurements were shown to identify stall and blade eigenmode frequencies, where flap bending bridges show a more reliable capture of stalled flow than torsional bridges. Furthermore, blade tip deflection measurements were shown to reduce with increased
This study examines the ability of a large (1200 lb gross weight) hexacopter with collective pitch controlled rotors to tolerate single motor failure. The hexacopter is considered in various orientations, and the vehicle is trimmed with one motor inoperative (OMI). Unlike RPM-controlled hexacopters, which were trimmable but uncontrollable in hover, and were untrimmable in cruise with an aft-rotor failure; with pitch-control the hexacopter is controllable in hover as well as trimmable for failure of any rotor in cruise (including an aft rotor failure). The study examines how pitch controls, and thrust are redistributed amongst the operational rotors, post-failure, for the different hexacopter orientations. For each case, the maximum thrust and torque increases on any individual rotor, and the total power increase, post-failure is examined. It is found that the hardest to trim cases are those where the hub torque and the hub drag induced yaw moment of the failed rotor add, and fault
The development of a coupled computational structural dynamics (CSD) and electrodynamic suspension (EDS) system was critical in modeling and predicting the aeromechanics of MagLev Aero's (MLA) propulsion system, ensuring safe testing and proving viability of levitated rotors for vertical lift systems. This advancement validates the feasibility of this enabling technology in applications of uncrewed aerial systems (UAS) with high hover lift efficiencies. This paper explores the implementation of an electromagnetic motor hub on a large-root-cutout, slowed rotor system with a specific focus on the impacts on aeromechanics: loads, performance, vibrations, and aeroelastic stability. The performance benefits of a large-root-cutout system, with an external or internal rotor, are well known; however, the mechanisms to implement such a design have been impractical. The development of an EDS motor bearing enables previously unattainable configurations like large-root-cutout and tip-driven ducted
The influence of ground, wall, and corner boundaries on multirotor vehicle performance was investigated through a series of controlled flight tests. Changes in rotor inflow profiles were represented by near-field rotor pressure measurements captured by a custom Kiel probe wake rake. Ground effect was characterized by reduced thrust and power requirements, primarily driven by the vehicle fuselage, which induced regions of reduced pressure and increased flow unsteadiness around the airframe. Operating near a wall boundary was found to restrict airflow into the portion of the rotor disk closest to the wall, leading to increased power requirements to maintain hover and a consequent reduction in performance. While vehicle orientation had minimal impact on overall rotor performance, it did influence local rotor inflow behavior near the wall, depending on the relative position of the interaction region formed with adjacent rotors. As the vehicle descends from the isolated wall effect into
During helicopter air-to-air refueling the rotor of the helicopter might enter the slipstream of the tanker aircraft's propeller. Based on blade element momentum theory, the impact of the accelerated air within the propeller slipstream on rotor blade aerodynamics (thrust, rolling and pitching moments) can be solved analytically. Also, DLR's comprehensive rotorcraft code has been used with the Pitt-Peters induced inflow plus rotor-rotor interference model. Additionally, DLR's free-wake code was used for both the propeller and the helicopter main rotor, including mutual wake-wake-interactions. The helicopter rotor's collective and cyclic controls needed for disturbance rejection are computed with all these models for a typical air-to-air refueling scenario without and with blade flapping motion. A propeller wake affecting the retreating side of the rotor requires much larger control inputs to retrim than an impingement on the advancing side. The results of all modelling approaches are
Rotor performance in a Martian environment was analyzed with an objective of increasing thrust with minimal impact on efficiency. The Sample Recovery Helicopter (SRH) and Rotorcraft Optimization for the Advancement of Mars Exploration (ROAMX) rotors were studied by varying solidity, blade count, and chord distribution to determine which configuration delivered the most desirable performance. For all configurations, the ROAMX rotor displayed better performance than the SRH rotor. It was observed that increasing solidity reduced the blade loading required to achieve the peak figure of merit, and beyond a solidity ratio of 0.3 the figure of merit was negatively impacted. For both rotors a 6-bladed configuration with a solidity ratio of 0.3 delivered the optimal figure of merit.
Inspecting the interiors of tanks and ships for defects involves accessing confined and elevated spaces. This can be difficult and hazardous for a person. Ducted aerial vehicles that can hover close to the object of interest can achieve this in a safer and more efficient manner. Such a vehicle is desired to be compact, to have a high hover endurance and to be protected from impact. This paper describes a design concept comprising ducted coaxial counter-rotating rotors with a compact swashplate mechanism for cyclic pitch input to the lower rotor. An experimental setup was used to investigate the effect of the duct. A numerical Blade Element Momentum Theory model was developed and validated to inform rotor selection. A prototype was designed and built with a hover thrust of 9.17 N, outer diameter of 350 mm, and height 173 mm. The duct provided a thrust benefit of 32% for this configuration for a given power. The prototype achieved stable controlled flight in hover and in passing near
Developed in the frame of the European Clean Sky 2 program, the RACER High Speed Helicopter Demonstrator of Airbus performed its maiden flight on April 25th, 2024. In the continuity of the previous high-speed demonstrator X3 (1st flight in 2010) the RACER is a 7/8t (15000 / 18000 lb) class compound helicopter powered by two SHE Aneto-1X engines, including a wing and two propellers. The tail rotor is removed as the two propellers control the yaw axis by differential thrust. At flight 07, with its initial default settings, it reached a true airspeed of 227 kts in level flight, exceeding its objective of 220 kts.
This paper explores a significant step forward, regarding the further detailed understanding of the Fenestron®. Since its patent in 1968 – for the Gazelle helicopter –, the shrouded tail rotor has been resized, inclined, modulated, etc. and has thus been continuously enhanced on different rotorcraft. Half a century after its invention, Airbus is once again exploring in more detail the magic of the Fenestron®, with the objective of optimizing it even further, for future helicopter applications. To grasp and observe properly some specific phenomena, a model (scaled to one third) capable of both unprecedented functions and modularities, was developed. The present paper will describe in detail the novel model and the related challenges and solutions. This model is capable of high rotor speed and dynamic pitch inputs, delivering power levels high enough to reach stall effects, while allowing the measurement of propulsive efficiency and to differentiate rotor vs fairing thrust. Furthermore
This study presents empirical modifications of Blade Element Momentum Theory (BEMT) to improve rotor performance prediction for open rotors in hovering conditions. The empirical adjustments were made to the inflow ratio, factoring in the real rotor wake area and estimated induced power losses. A comparison between experimental data and two analytical models, one using an empirical inflow formula and the other a theoretical formula (classical BEMT), was conducted for two rotors. The empirical inflow model demonstrated superior accuracy in predicting thrust and torque. These modifications are applied to the inflow ratio by accounting for the actual rotor wake area and estimated induced power losses. The findings highlight the potential for more accurate performance prediction through the integration of empirical data into theoretical frameworks.
The experimental investigation analyzed the performance of three machining conditions: dry machining, cryogenic machining, and cryogenic machining with minimum quantity lubrication (MQL) on tool wear, cutting forces, material removal rate, and microhardness. The outcome of this study presents valuable knowledge regarding optimizing conditions of turning operations for Ti6Al4V and understanding the machinability under cryogenic-based cooling strategies. Based on the experimentation, cryogenic machining with MQL is the most beneficial approach, as it reduces cutting force and flank wear with a required material removal rate. This strategy significantly enhances the machining efficiency and quality of Ti6Al4V under variable feed rates (0.05 mm/rev, 0.1 mm/rev, 0.15 mm/rev, 0.2 mm/rev, 0.25 mm/rev) where cutting velocity (120 m/min) and depth of cut (1 mm) are constant. The effects of the main cutting force, feed force, thrust force, material removal mechanism, flank wear, and
NWI Aerostructures Nashville, TN
In any human space flight program, safety of the crew is of utmost priority. In case of exigency in atmospheric flight, the crew is safely and quickly rescued from the launch vehicle using Crew Escape System (CES). CES is a critical part of the Human Space Flight which carries the crew module away from the ascending launch vehicle by firing its rocket motors (Pitch Motor (PM), Low altitude Escape Motor (LEM) and High altitude Escape Motor (HEM)). The structural loads experienced by the CES during the mission abort are severe as the propulsive, aerodynamic and inertial forces on the vehicle are significantly high. Since the mission abort can occur at anytime during the ascent phase of the launch vehicle, trajectory profiles are generated for abort at every one second interval of ascent flight period considering several combinations of dispersions on various propulsive parameters of abort motors and aero parameters. Depending on the time of abort, the ignition delay of PM, LEM and HEM
A new framework for performing high-fidelity computational aeromechanics simulations of the V-22 tiltrotor aircraft in hover mode has been developed. It is built on the HPCMP CREATE-AV Helios tool and utilizes scripted input generation and automatic replacement of modular model components. This new framework has been used to investigate the impact of various approaches to modeling the rotor aerodynamics, airframe aerodynamics, and periodic blade motion on predictions of aircraft hover performance in and out of ground effect. The findings indicate that an actuator line method can provide rotor performance predictions with comparable accuracy to a meshed-blade approach. However, body-fitted meshes are required to compute accurate airframe download. Furthermore, active trimming of rotor collective to a target thrust provides more representative aircraft aerodynamic performance than directly applying a collective angle as measured during flight test. The computational framework can
This study models the interaction of a two-bladed 14" propeller with the ground under different configurations using actuator disk method (ADM) where the rotor is modeled using unsteady momentum sources distributed over the entire disk. While ADM has been extensively used for standard rotorcraft analysis, it's performance in unconventional operating conditions remains an open question. Exhaustive experiments conducted at DEVCOM Army Research Laboratory are compared with ADM to evaluate the inexpensive method's ability to predict rotor loads for parametric variations in rotor-ground interaction scenarios. Partial ground effect (part of the rotor operating IGE), side-by-side rotors in ground effect and variation in IGE pitch attitude are specifically considered in this study. ADM generally predicts the thrust increase in partial ground effect (PGE) as the rotor goes from OGE to IGE although the increase is somewhat earlier and milder than measured in experiments. Side-by-side rotors in
The effectiveness of the counter-torque capabilities of rotorcraft tail rotors has undergone extensive study and development due to its critical role in balance and maneuverability. The Fenestron provides an interesting alternative to the conventional open tail rotor, improving upon several key elements, but still demonstrates drawbacks in others. This paper seeks to provide a side-by-side comparison of Fenestron and open tail rotor performance, in terms of thrust and power, using the mid-fidelity surface-vorticity panel-method flow solver FlightStream® to simulate Fenestron and open rotor configurations in hover and cruise conditions. While this work is not a comprehensive analysis, as it does not compare to experimental data, nor implements exact geometrical representation, it does propose trends between the Fenestron and open tail rotor cases.
Multicopters operate in environments subject to strongly gusting winds, and need good aeromechanical models to improve the aircraft. A common, convenient, assumption is that the gusting inflow is quasi-static at each instant, but this assumption has never been tested. This paper shows that there is a solid physical basis for the simplified aerodynamic models of multicopter response to gusts. Experiments and computations show that using the static relationship between thrust or power and aerodynamic angle of attack for a multicopter rotor (the quasi-static assumption) in sinusoidally pitching sideflow can be used to predict the thrust or power for unsteady variation of angle of attack if the instantaneous flow angle of the freestream is known. Vertical (angle) gusts up to 1885°/s (k=2.2 based on diameter) and with a wavelength longer than the rotor diameter were shown to be covered by this assumption.
Rotorcraft responses to idealized disturbances are examined to gain insights into model fidelity requirements for flight simulations of the ship-rotorcraft dynamic interface. Two disturbance fields are considered: an isolated straight vortex that represents the canonical vortex that results from the corners of flat top ships in oblique wind-over-deck conditions and a horseshoe vortex derived from a nondimensional characterization of the time-averaged flow observed aft of a simplified ship superstructure. Rotorcraft models considered include: an analytical blade element theory-based rotor model, where the disturbance velocities are integrated over the rotor, and a coupled blade element / free wake flight dynamic model of the full UH-60 aircraft, which is used to perform time-marching simulations with the disturbances modeled as a frozen field that is fixed in space and not interacting with the aircraft (one-way coupling), and as a distorting field (two-way coupling). Analytical thrust
This study uses a mid-fidelity, aeromechanics coupled framework using the Lattice-Boltzmann Method (LBM) fluid solver to investigate an experimental coaxial rotor system. Co-rotating and counter-rotating rotor operation scenarios in hover are studied. The rotors are represented as an actuator line in the LBM fluid simulation. Simulation results are compared to experimental data, mid-fidelity and CFD simulation results available in literature. Results indicate that the framework can accurately predict thrust and thrust variations for both upper and lower rotors. Power prediction has deficiencies compared to experimental and CFD results, but is in line with mid-fidelity simulation results in literature. Flow field results are also compared qualitatively with CFD results. Results are sensitive to the actuator line representation of the blade, the inflow sampling, and tip corrections.
This study investigates the use of machine learning (ML) models to estimate the gross weight (GW), the longitudinal position of the center of gravity (CGx), and 1/rev cyclic flapping angles (Δ1c and Δ1s) of a compound helicopter with three redundant controls - main rotor RPM, collective propeller thrust, and stabilator angle. Neural Network (NN), Gaussian Process for Regression (GPR), and Support Vector Machine (SVM) algorithms are employed to develop estimation models using supervised training. The airspeed, redundant controls, main rotor controls, aircraft attitudes, and main rotor torque are selected as input variables (predictors) to the models due to their accessibility through the aircraft Health and Usage Monitoring System (HUMS). The dataset is split into low-speed and high-speed regimes to compare the prediction accuracy and training cost of separate regime models against a combined full-regime model. Separate airspeed regime GPR models showed superior performance in GW
This study models flow around isolated and side-by-side three-bladed propellers in (IGE) and out of ground effect (OGE) using actuator-based techniques of varying fidelity. Actuator techniques model propellers using momentum sources distributed over the disk in actuator disk method (ADM) or distributed over moving lines in actuator line method (ALM) to reduce computational cost compared to blade-resolved DDES simulations. The lowest fidelity ADM method is observed to reasonably predict thrust with the use of a tip loss model to control runaway thrust at the tip while not resolving flow features such as blade-bound vortices and helical tip vortices at a fraction of the cost of BR-DDES (1/100). The coarser ALM model resolves these features but still requires a tip loss model to control runaway thrust at 1/10th the cost of BR-DDES. Finally, the finer ALM model used in this study accurately captures blade-related features and further predicts the tip loss trend from first principles at 1
This paper addresses the urgent need to enhance rotorcraft safety and performance by developing a prediction methodology for the onset of the Vortex Ring State (VRS), and therefore verifying the VRS avoidance diagram. The objectives of this research are to assess the correlation between predictions generated by a comprehensive flight dynamics code and the latest and most accurate VRS boundary models, validate the VRS avoidance diagram across diverse descending flight conditions, and identify specific parameters indicating the rotor's entry into the VRS. The methodology involves a detailed investigation of 8 descent manoeuvres using a comprehensive flight dynamics code coupled with an advanced free vortex wake model. Results show that the pitch and roll oscillations and thrust fluctuations experienced by helicopters during the VRS are also observed in the model response to steep descent maneuvers. The findings confirm the reliability and applicability of the VRS avoidance diagram
Airfoil optimization for rotor blades is a critical endeavor aimed at enhancing aerodynamic performance and reducing noise. This paper employs a Kriging surrogate model combined with a multi-objective genetic algorithm to optimize thrust, power, and broadband noise. Three airfoil parameterization methods including ParFoil, PARSEC, and CST are compared when used to generate various airfoil shapes for the surrogate model and optimization process. We utilize low-fidelity aerodynamic tools such as XFOIL and blade element momentum theory for aerodynamics. In addition, acoustic modeling is conducted using Lee's wall pressure spectrum model alongside Amiet's trailing-edge noise model. The paper focuses on small-scale rotor configurations, specifically an ideally twisted rotor using the NACA 0012 airfoil and a modified XV-15 blade. Both blades are used as baseline models for hover optimization. The optimization of the ideally twisted rotor across various parameterization methods demonstrates a
A quadrotor was modified by adding wings to the frame to directly compare the flight dynamics characteristics as well as the stability and control derivatives of the quadrotor and its biplane tailsitter variant. The on axis response of the quadrotor and a biplane tailsitter variant were measured through flight test and frequency domain system identification was used for non-parametric and parametric model identification. Identification of the full vehicle dynamics demonstrated that also identifying the motor torque and back-EMF constants from no-load measurements and the remaining motor parameters from a rotor-motor test stand provided the most accurate identified full vehicle model. The motor dynamics were shown to add a pole to the thrust-based responses (roll, pitch, and heave), while the torque based response (yaw) included a pole and a zero. This approach was then used to identify and compare the quadrotor dynamics, tailsitter dynamics, and the total impact of canting the motors
The Shake-The-Box technique was applied to experimentally quantify the time-resolved volumetric flow field around a free-flying quadcopter UAV with an overall span of about 0.5 m. State-of-the-art LED illumination and high-speed camera equipment was combined with modern Lagrangian tracer particle tracking and data assimilation techniques, facilitating a measurement volume larger than 1.5m3. The setup allowed for both hover and limited maneuvering of the quadcopter, while resolving even small details of the complex interactional aerodynamics. In hover out of ground effect, the four individual rotor wakes merged into a single jet within a few rotor radii below the rotor planes. Evaluating the mass and momentum fluxes over suitable control volumes yields accurate estimates for the quadcopter's total thrust, the asymmetric thrust distribution between front and back rotors, and the entrainment of external flow through turbulent mixing. Hover in ground effect decreases the power requirement
The flow behavior of the two-blade MERIT rotor in hover, focusing on both pre-stall and stall regimes, is investigated through a comprehensive numerical-experimental approach. The study leverages unsteady RANS simulations to compute rotor thrust and power polars and validates them against experimental measurements. Valuable insights are provided into the capabilities of unsteady RANS methods and modern turbulence models for predicting rotor performance across these critical operating conditions. Furthermore, the numerical model incorporates blade deformations by implementing the experimentally measured flap and torsion displacements. A more realistic depiction of the rotor's aerodynamics is provided accounting for the structural deformations of the blades under aerodynamic loads. Highfidelity simulations closely predict the experiments in pre-stall conditions while discrepancies are present when the flow exhibits extended stalled regions. Blade deformations demonstrated to have only a
Installation effects of the Volocopter 2-X beam structures are studied by performing high-fidelity CFD simulations of a single and three-rotor configurations in hover. The studied cases are compared with simulations without airframe to investigate the installation effects. In addition, the noise emission of the configurations is simulated by using a Ffowcs Williams-Hawkings based CAA code. Scattering effects are also included by using a BEM code. The rotors are simulated at an identical RPM and are placed in their mounting position. Furthermore, an additional setup with individual rotor RPMs is simulated for the three-rotor configuration. The installation mainly affects the rotor wake, thrust and pressure fluctuations on the rotor, while the integral aerodynamic quantities remain almost unchanged. This resulted in additional oscillations in the acoustic pressure signal. Overall, the installation increases the OSPL by about 1.5 dB, but has a greater effect on the 3-20 harmonics. The
A towing tank investigation of a single rotor blade operating at hovering and high advance ratio conditions is presented. A custom blade was manufactured and instrumented with fully bridged axial strain gauges to monitor the flap bending strain at three radial locations. Measurements of rotor thrust and torque were obtained to characterise the rotor aerodynamic environment for advance ratios ranging from 0.4 to 1.00 and to identify the presence of stalled and reverse flow. Strain measurements obtained at three locations across the blade span show minima and maxima at approximately the same azimuthal location as the load data. Moreover, the strain distribution shows a growth in strain magnitude with increasing advance ratio. Spectra of strain shows a dominant 1/rev signal and for the ∅ = 25° collective, non-harmonic frequencies are observed due to aperiodic vortex shedding from the presence of stalled flow.
Thrust measurement systems come in many sizes and shapes, with varying degrees of complexity, accuracy and cost . For the purposes of this information report, the discussions of thrust measurement will be limited to axial thrust in single-axis test systems.
In an application first, the physics of why the sky is blue is used to measure gas flows without obstructive sensors. A longstanding industry partnership between Virginia Polytechnic Institute and State University (Virginia Tech) and Pratt & Whitney has resulted in a new laser-optical technology that aims to revolutionize in-flight thrust measurement.
This document defines and illustrates the process for determination of uncertainty of turbofan and turbojet engine in-flight thrust and other measured in-flight performance parameters. The reasons for requiring this information, as specified in the E-33 Charter, are: determination of high confidence aircraft drag; problem rectification if performance is low; interpolation of measured thrust and aircraft drag over a range of flight conditions by validation and development of high confidence analytical methods; establishment of a baseline for future engine modifications. This document describes systematic and random measurement uncertainties and methods for propagating the uncertainties to the more complicated parameter, in-flight thrust. Methods for combining the uncertainties to obtain given confidence levels are also addressed. Although the primary focus of the document is in-flight thrust, the statistical methods described are applicable to any measurement process. The E-33 Committee
Items per page:
50
1 – 50 of 806