Thrust Reduction Research Articles

PLR and take-off speed implications on noise contours and local gaseous emissions for SST business jet concept aircraft

DLR has established a simulation process to evaluate the landing and take-off of novel civil supersonic transport (SST) aircraft. This process was applied to two business jet concept aircraft in order to assess the certification noise levels according to novel regulations under discussion at ICAO CAEP. It has to be noted, that neither of the concept aircraft is optimized for low noise nor for low boom. At DLR, modifications to departure flight procedures are under investigation, i.e., an automated thrust reduction referred to as Programmed Lapse Rate and an increased take-off speed. These measures as proposed by NASA promise to reduce the certification noise levels for the two departure locations. For the two considered concept aircraft, promising procedures with reduced noise at the certification locations have been identified previously. This paper now addresses the impact of these promising departure procedures on noise contour areas and local gaseous emissions. Impact on EPNL and SEL contours and differences in CO2 and NOX are assessed along the simulated flights. It is demonstrated, that the proposed procedures do not only reduce noise at the certification locations but also noise contour areas and gaseous pollutants.

Numerical study on flow field characteristics and performance of a Mach 10 internal injection kerosene-fueled oblique detonation engine

Oblique detonation engines (ODE) have significant potential for hypersonic propulsion, yet there is a paucity of research investigating internal injection ODEs. In this study, a numerical simulation of the internal flow field of a Mach 10 internal injection kerosene-fueled ODE is conducted. The fuel mixing, pre-combustion, and combustor wave structure in the flow field are analyzed in a situation closer to the real flow field conditions. Further studies have demonstrated that alterations to the upper wall initiation position of nozzle can influence the separation zone, flow field stability and the engine performance. An upper wall initiation position that is too far forward will increase the separation zone area and reduce the engine thrust. Conversely, an upper wall initiation position that is too far back will lead to flow field destabilization and eventually thermal choking. Finally, the effects of increasing the equivalent ratio on the flow field structure and engine performance for a certain configuration are analyzed. The results demonstrate that when the equivalent ratio is elevated, an increase in either the bottom or top incoming flow equivalent ratio results in a transformation of the wave structure within the combustor due to the presence of the incoming boundary layer and the subsonic zone. A large-scale separation zone will form at the bottom of the combustor, resulting in a reduction in nozzle thrust. However, the wedge drag is reduced more, thereby increasing the engine's specific impulse.

Optimizing the Landing Stability of Blended-Wing-Body Aircraft with Distributed Electric Boundary-Layer Ingestion Propulsors through a Novel Thrust Control Configuration

The imperative for energy conservation and environmental protection has led to the development of innovative aircraft designs. This study explored a novel thrust control configuration for blended-wing-body (BWB) aircraft with distributed electric boundary-layer ingestion (BLI) propulsors, addressing the issues of sagging and altitude loss during landing. The research focused on a small-scale BWB demonstrator equipped with six BLI fans, each with a 90 mm diameter. Various thrust control configurations were evaluated to achieve significant thrust reduction while maintaining lift, including dual-layer sleeve, separate flap-type, single-stage linkage flap-type, and dual-stage linkage flap-type configurations. The separate flap-type configuration was tested through ground experiments. Control experiments were conducted under three different experimental conditions as follows: deflection of the upper cascades only, deflection of the lower cascades only, and symmetrical deflection of both cascades. For each condition, the deflection angles tested were 0°, 10°, 20°, 30°, 40°, 50°, and 60°. The thrust reductions observed for these three conditions were 0%, 37.5%, and 27.5% of the maximum thrust, respectively, without additional changes in the pitch moment. A combined thrust adjustment method maintaining a zero pitch moment demonstrated a linear thrust reduction to 20% of its initial value. The experiment concluded that the novel thrust control configuration effectively adjusted thrust without altering the BLI fans’ rotation speed, solving the coupled lift–thrust problem and enhancing BWB landing stability.

Thrust Generation Through a Collimated Electron Beam in a Magnetic Mirror and its Possible Applications in Plasma Thrusters

In this article, we apply some basic concepts of electric thrusters, using Monte Carlo simulations. One such concept is the thrust, which we calculate for the simulated case of a simple thruster, consisting of two stages. The goal is to study the thrust in a magnetic bottle as a function of the collimation of an assembly of particles. In the first stage, a voltage applied between a pair of electrodes accelerates the particles. We perform simulations of the behavior of the assembly of charged particles. A normal distribution is used for the initial velocities of the assembly, with a given standard deviation. The resulting velocity distribution and its dependence on the applied voltage are analyzed. Further, the set of particles, whose pitch angles follow a normal distribution with a given standard deviation, enters a magnetic bottle where it is collimated to reduce its dispersion upon the exit of the thruster. In this second stage, we study the conditions for which the collimation produced, by the magnetic bottle, leads to the least thrust reduction. For the simulations, we assume electrons; However, the concepts apply to any charged particle species. The simulations show that the standard deviation of the pitch angle distribution has an important influence on the confinement of charged particles and on the thrust that can be generated when collimating the particle beam in a magnetic bottle.

Experimental Analysis of Robotic Cortical Bone Specimen Drilling Performance: Effect of Cryogen.

In orthopedic surgery, precise bone screw insertion is crucial for stabilizing fractures, necessitating a preliminary cortical bone drilling procedure. However, this process can induce temperatures exceeding 70 °C due to the low thermal conductivity of cortical bone, potentially leading to thermal osteonecrosis. Furthermore, significant cutting forces and torque pose risks of tool breakage and bone damage, underlining the need for high precision and optimal processing parameters. Traditionally, drilling relies on the surgeon's experience and often results in imprecise outcomes due to inconsistent feed rates. Therefore, this study proposes the use of a 6-axis robot for controlled drilling, offering precise control over angular velocities and consistent feed rates. Additionally, explore the use of cryogenic liquid nitrogen (LN2) as a novel cooling method compared to conventional saline solutions, examining its efficacy under various cutting conditions. The results demonstrate that LN2 cooling conditions lead to a reduction in thrust and torque under specific processing conditions, and facilitate smoother chip evacuation. Additionally, LN2 significantly lowers the peak temperature around the drilling site, thus minimizing the risk of thermal osteonecrosis. Consequently, the use of a 6-axis robot provides consistent feed rates, and LN2 cooling achieves optimal processing conditions, enabling a more controlled and effective drilling process.

Wake interference of tandem wind turbines considering pitch strategy based on the AL-LDS-Ωnew coupling method

This paper establishes a high-fidelity and efficient Computational Fluid Dynamics (CFD) numerical method (AL-LDS-Ωnew) for wind turbine wake by combining the actuator line (AL), the localized dynamic Smagorinsky (LDS) sub-grid scale (SGS), and the new generation Ωnew vortex identification method under the framework of large eddy simulation. The model advantages are encouraging: 1) In terms of turbine modeling, the AL model is adopted to replace the traditional three-dimensional solid model, which avoids solving the boundary layer on the blade surface and improves computational efficiency; 2) In terms of wake simulation, the LDS SGS model is used to model turbulence, reducing vortex dissipation and further improving the refinement of turbine wake; 3) In terms of vortex identification, the new generation Ωnew vortex identification method avoids the difficult threshold selection in previous vortex identification and captures more refined vortex structures. The accuracy of the model is validated against published data of a NREL 5 MW wind turbine, and then extended to simulate the wake interference of tandem twin-rotor turbines by changing the pitch angle of the upstream wind turbine (WT1). The influence mechanisms between array wake interference and energy conversion efficiency under the pitch strategy are explored, demonstrating the AL-LDS-Ωnew coupling method is computationally accurate and efficient for simulating the complex wake interference. From analyses, the pitch strategy can effectively suppress the wake effect of the upstream turbine (WT1) and increase the power output of the downstream turbine (WT2), thus improving the overall output power of the array farm. Compared with the non-pitch condition (0 pitch angle), a pitch angle of (2°) maximizes the global energy conversion efficiency of the twin-rotor array: power augmentation by 0.29%, and thrust reduction by 5%. This optimal state reduces the fatigue load of the turbine and is more conducive to long-term operation. The findings, whilst preliminary, encourage the use of turbine pitch strategies in the wind farm planning and operation.

Study on Aerodynamic Performance and Wake Characteristics of a Floating Offshore Wind Turbine in Wind–Wave Coupling Field

Floating offshore wind turbines (FOWTs) exhibit complex motion with multiple degrees of freedom due to the interaction of wind and waves. The aerodynamic performance and wake characteristics of these turbines are highly intricate and challenging to accurately capture. In this study, dynamic fluid body interaction (DFBI) and overset grid technology are employed to investigate the dynamic motion of a 5 MW FOWT. We use the volume of fluid (VOF) method and improved delayed detached eddy simulation (IDDES) model to investigate the aerodynamic performance and wake evolution mechanism for various wave periods and heights. According to the findings, the magnitude of the pitch motion increases with the period and height of the waves, leading to a decrease in both the power output and thrust; the maximum power was reduced by nearly 6.8% compared to a wind turbine without motion. The value of power and thrust reduction varies for different wave periods and heights, and is influenced by the relative speed and pitch angle, which play a crucial role. Wind–wave coupling has a significant impact on the evolution of both wake and vortex structures for FOWT. The wake shape downstream is also dynamically influenced by the waves. In the presence of wind and wave coupling, the interaction between the wind turbine and the wake is heightened, leading to the merger of two unstable vortex rings into a single, larger vortex ring. The research unveils a comprehensive picture of the offshore wind energy dynamics and wake field, which holds immense significance for the design of floating wind turbines and the optimization of wind farm layout.

Numerical Investigation of Seismic Behaviour for a Flexible Cantilever Retaining Wall with Cohesive Backfill

In the seismic design of flexible cantilever walls retaining cohesive backfill soil, the common practice is to neglect the cohesion effect. Dynamic lateral earth pressure is typically evaluated based on approaches primarily intended for cohesionless soils or through analytical pseudo-static methods. Nevertheless, both experimental and theoretical evidence has demonstrated significant effects due to soil cohesion that are not accounted for by these methods. This study involved finite element modeling (FE) of a flexible cantilever wall with a height of 5.4m, supporting homogeneous cohesive backfill under initial static and seismic loadings. The calculated active earth thrust was then compared with values obtained experimentally and through conventional analytical methods. The obtained results indicate that the presence of soil cohesion significantly reduces seismic demands on flexible cantilever retaining walls, resulting in a substantial reduction of seismic active earth pressures and total seismic thrust by up to 50% and 52%, respectively. It enhances also the overall stability of the system by shifting the point of application of seismic thrust toward the base of the wall, thereby increasing the safety margin. In addition, it significantly decreases the wall displacement at the stem top, with reductions of up to 104% compared with the case involving cohesionless backfill. It was observed that the conventional methods recommended by some seismic regulations largely underestimate seismic active pressure.

Design, Modeling, and Control of a Composite Tilt-Rotor Unmanned Aerial Vehicle

Tilt-rotor unmanned aerial vehicles combine the advantages of multirotor and fixed-wing aircraft, offering features like rapid takeoff and landing, extended endurance, and wide flight conditions. This article provides a summary of the design, modeling, and control of a composite tilt-rotor. During modeling process, aerodynamic modeling was performed on the tilting and non-tilting parts based on the subcomponent modeling method, and CFD simulation analysis was conducted on the entire unmanned aerial vehicle to obtain its accurate aerodynamic characteristics. In the process of modeling the motor propeller, the reduction of motor thrust and torque due to forward flow and tilt angle velocity is thoroughly examined, which is usually ignored in most tilt UAV propeller models. In the controller design, this paper proposes a fusion ADRC control strategy suitable for vertical takeoff and landing of this type of tiltrotor. The control system framework is built using Simulink, and the control algorithm’s efficiency has been verified through simulation testing. Through the proposed control scheme, it is possible for the composite tiltrotor unmanned aerial vehicle to smoothly transition between multirotor and fixed-wing flight modes.

Mini Gaz Türbini (MGT) Motorunda Aspir Metil Ester-Jet A1 Karışımlarının Kullanımının Performans ve Emisyonlara Etkisinin Araştırılması

This study focused on the investigation of the effect of the SME-Jet A1 mixture on the thrust performance, fuel consumption and pollutant emissions of a small-scale jet engine, also the alternative of safflower methyl ester (SME) to Jet A1 fuel. The experiments were carried out by using bio jet fuels obtained by mixing Jet A1 and the fuel produced from safflower oil in taxi, approach, climb and take-off power cycles of a jet engine. The thrust forces of both types of fuel obtained from the percentage power values of the aircraft under real operating conditions were determined and accordingly CO, HC, CO2 emissions, fuel consumption and exhaust gas temperature measurements were made. As a result of the measurements, approximately 27.5% reduction in thrust was observed with the use of biojet. In addition, HC emissions decreased by up to 51%, while CO emissions increased by 30% at take off while being closer to each other at low rpm. While the CO2 emissions of methyl ester increased by 8-16% depending on the speed change, fuel consumption, while at an acceptable level at low rpm, exceeded 50% at take off.

Continuous Injector Geometry Variation to Augment Rotating Detonation Combustor Operation and Performance

Sensitivity of rotating detonation combustor operation and performance to oxidizer injector pressure drop was characterized using continuous variation of the injector area during combustor operation. As the oxidizer injector area was both increased and decreased, the sensitivity of the combustion process to varying injector pressure drop was characterized using high-frequency measurements of pressure and chemiluminescence intensity. Detonation wave strength and coherence were characterized using peak-to-peak intensity and power fraction calculated from point-chemiluminescence measurements. Propulsive performance of the combustor was evaluated using thrust and equivalent available pressure, relating them back to reactant supply pressures for assessment of combustor pressure gain. Pressure gain increased during a test as the oxidizer injector area was increased and the corresponding manifold pressure was decreased. At larger injector areas, pressure gain decreased as the operating mode of the combustor transitioned from detonation to deflagration, concomitant with a reduction of gross thrust. Modeling of injector recovery time revealed that the injector operated in both choked and unchoked regimes, which was used to explain detonation wave number transitions in the experiment. A broadened range of detonative operability enabled by active variation of combustor geometry resulted in higher performance with a lower injector pressure drop.

Experimental and numerical investigation of the propeller thrust reduction in bollard pull condition

The polar ships are often in bollard pull condition, which is prone to induce propeller cavitation and cause thrust reduction, weakening the ship’s ice-breaking ability. Based on the model tests and numerical simulations, this paper aims to investigate the propeller thrust reduction in the bollard pull condition. Firstly, the test of a self-designed ice class propeller under different cavitation numbers is carried out. The thrust reduction phenomenon is successfully captured and the hydrodynamic characteristics are obtained. Subsequently, the simulation based on Star-CCM+ is performed. It is found that the propeller thrust decreases almost linearly after the cavitation number reaches a certain threshold. When the cavitation number is equal to 0.25, the thrust decreases to 38.1% of that in the open water condition. By analyzing the mechanism of thrust reduction, the main reason is recognized that the average pressure of the blade face decreases greatly due to the influence of the cavitation on the blade back.

Fully kinetic study of facility pressure effects on RF-source magnetic nozzles

A fully kinetic 2D axisymmetric Particle-in-Cell (PIC) model is used to examine the effects of background facility pressure on the plasma transport and propulsive efficiency of magnetic nozzles. Simulations are performed for a low-power (150 W class) cathode-less radio-frequency (RF) plasma thruster, operating with xenon, between background pressures up to 10−2 Pa and average electron discharge temperatures of 4–16 eV. When the electron temperature within the near-plume region reaches 8 eV, a decisive reduction in performance occurs: at 10−2 Pa, in-plume power losses surpass 25% of the discharge energy flux. Given that the ionisation energy for Xe is 12 eV, the 8 eV threshold indicates that a consistent percentage of electrons has energy enough to trigger ionisation. On the other hand, when the temperature is below such threshold, the primary collisions are charge-exchange and inelastic ion scattering, and the power loss remains less than 10%. It is established that losses in the considered thruster are significant if the facility pressure is greater than 10−3 Pa, at absorbed powers larger than 130 W. At the nominal 150 W, this results in a 15% thrust reduction. When facility pressure is taken into consideration over ideal vacuum simulations, numerical error is reduced to <30% when compared to experimental thrust measurements at 10−3 Pa.

An investigation of the noise characteristics of a small-scale rotor in axial descent

In this work, the acoustic characteristics of a small-scale drone rotor in axial descent are numerically investigated. Various descent speeds covering different working states are tested, and the corresponding effects on the noise are analysed. The turbulent flow fields are obtained using delayed detached eddy simulations and an acoustic-wave preserved artificial compressibility method. The far-field noise is obtained by an integral solution of the Ffowcs-Williams and Hawkings equation. The aerodynamic force and the noise features are also evaluated experimentally in anechoic wind tunnel tests, and test data are used to validate the numerical simulations. The results show that the increase in descent speed leads to a reduction in the mean thrust. Significant thrust fluctuations are also observed when the rotor approaches the vortex ring state. Consequently, a substantial increase in sound pressure level in the low to medium frequency range is found in the far-field noise spectra, while the high-frequency sound pressure level shows a slight decrease. Furthermore, a noise source analysis is performed on the blade surface to help understand the corresponding noise mechanism.

Conceptual design of a pilot assistance system for customised noise abatement departure procedures

The departure of an aircraft is commonly the flight phase with the highest thrust level in a flight, which leads to considerable noise levels on the ground. The departure procedures are characterised by the thrust reduction altitude, the acceleration altitude, and the control of airspeed and aircraft configuration within each take-off segment. Thrust reduction and acceleration altitudes are typically constant and are not adjusted to particular operational key parameters (e.g., take-off mass, reduced take-off thrust) or weather conditions. However, the parameters differ between individual flights and affect flight performance as well as the noise levels on the ground. This paper presents the conceptual design of a pilot assistance system which aims to reduce the noise on the ground by identifying a custom thrust reduction and acceleration altitude for existing noise abatement departure procedures. The pilot assistance system is based on an aircraft simulation model and a noise simulation and is aimed to be utilised during pre-flight planning on ground. A key part of the noise evaluation is that the local population distribution around the airport is considered. An overview of the ongoing research at the German Aerospace Center is provided. The conceptual design and preliminary results are presented and discussed using an exemplary take-off for three wind conditions.

Numerical investigations on aero-engine icing characteristics at mixed phase conditions

Ice crystal/mixed phase icing in aero-engine will cause engine surge, stalling, thrust reduction, and even stop in extreme cases, which seriously endangers flight safety. In this paper, the characteristics of mixed phase ice accretion in aero-engines are numerically studied. The air flow field model, mixed phase trajectory model, impact model, sticking model and ice accretion model are established. Meanwhile, the sensitivity parameters of ice accretion, such as flight speed, icing time, melting rate (LWC/TWC), and relative humidity (RH), were investigated in detail. The results show that the ice thickness of nacelle surface and nosecone surface increases with the increase of flight speed. With the increase of fan span, the icing thickness of the fan surface increases significantly. The icing of the fan is mainly concentrated on the pressure side of the fan, while the suction side of the fan only has a small amount of icing in the leading edge. In the early stage of icing, the icing thickness of nacelle increases rapidly, and with the increase of icing time, the icing thickness increases, but the icing growth rate becomes slower. With the increase of the melting rate, the ice thickness of the nacelle, the nosecone and the fan pressure side surface increased significantly. When the fan span is 90 % and the melting rate is 0.05, 0.15, and 0.25, the maximum icing amount on the fan surface is 1.56, 3.85, and 5.71 mm, respectively. When the RH is high, there is more ice accumulation on the nacelle surface. At RH is 80 %, the ice thickness at the nosecone tip reaches 5.48 mm. Finally, when using the hot air deicing method, the hot air flow needs to be reasonably distributed according to the amount of icing, and the parts with less icing can appropriately reduce the distribution of hot air for rational use of energy and energy conservation.

A novel in-flight thrust modulation technique for SPRM: proof of concept

Thrust modulation (termination or reduction) is one of the most sophisticated topics in rocket motors. Thrust termination is employed when combustion must be terminated to ensure proper stage separation, to avoid motor explosion, to attain certain range, or for research purposes. While liquid and hybrid propellant rocket motors have the ability to extinguish thrust, thrust termination in solid propellant rocket motors (SPRMs) need the use of specialized technology. In-flight thrust termination of (SPRMs) can be achieved by sudden and sufficient increase in throat area. Increasing throat area causes high depressurization rate which consequently leads to extinguishing the rocket motor. In contrast, in-flight thrust reduction can be theoretically attained either by reducing chamber pressure to a new equilibrium level or by reducing the divergent section length. The objective of the present paper is to prove the concept of using Linear Shaped Charge (LSC) for thrust modulation; a technique novel to SPRM applications. Thrust termination concept is proved via a set of static firing tests on a standard test motor. In addition, a mathematical model is developed to predict the drop in thrust due to cutting the nozzle at any arbitrary location. The model is validated by comparing with published experiments for thrust termination and reduction. A parametric study is conducted based on model results. It is confirmed that by cutting the nozzle along its divergent section, thrust reduction less than 20% can be attained. Further reduction can be achieved by cutting the nozzle along its convergent section.

Rolling-Sliding Performance of Radial and Offset Roller Followers in Hydraulic Drivetrains for Large Scale Applications: A Comparative Study

Generally speaking, excessive side thrust and roller slippage are two different aspects affecting cam-roller mechanisms. In novel large-scale hydraulic drivetrains for offshore wind turbines, the highly dynamic nature of these mechanisms combined with the interplay of cyclic loads, frictional torques and inertia promote slippage at the cam-roller interface. At larger scales, the effects of roller inertia become much more pronounced, as the inertia escalates exponentially with the roller’s radius. This study presents a comparative analysis between radial and offset roller followers in novel large-scale hydraulic drivetrains, where offset followers are incorporated to minimize the side thrust. The framework encompasses a comprehensive kinematic and force analysis, to provide the inputs for two lubrication models integrated into the torque-balance equation, where the possibility of slippage is allowed. The findings reveal that the equivalent side thrust can be reduced by 51% with offset followers. Both configurations experience slippage during the low-load phase, but it rapidly diminishes during the high-load phase. This sudden transition in rolling conditions results in a sharp increase in surface temperature and traction force, emphasizing the importance of minimizing sliding at the interface. However, besides the substantial side thrust reduction, offset followers showed superior tribological performance, mitigating undesirable thermal and frictional effects.

Monitoring of friction stirred spot weld quality for dissimilar Al6061 to polycarbonate using tool assisted thrust-torque signatures

Friction stir processing procedures are often found to be an alternative choice for the joining of thermoplastic polymers without melting the base materials. In recent years, dissimilar metallic to thermoplastic assemblies are highly used in aerospace and automotive industries as mechanical fastener or fusion welding techniques are rather inadequate for extreme discrepancy in thermal properties. The present work addresses the feasibility of friction stirred spot welding for non-ferrous aluminium (Al 6061) to polycarbonate using taper cylindrical tool of hardened H13 steel along with machine integrated axial thrust and stirring torque signals. The real-time positional and axial velocity of the rotating tool has been acquired to distinguish between plunging and dwelling stage. The process stability that is, the deviation of thrust-torque in respective phases was processed. The interaction of plunge depth and dwell time on tool pin assisted keyhole with associated material expulsion for different tool revolving speed was established. The weld mechanical behaviour has been scanned using load versus elongation diagram with associated weld pair interface. Lastly, sensors’ based process monitoring extended to weld quality prediction have been made using various significant features of thrust-torque-velocity based approaches. The tool spinning speed followed by dwell time was found to be important on weld quality index. The joint efficiency was significantly upgraded (more than 60%) at intermediate dwell time (40 s) with low tool plunging (0.3 mm) or reversed parametric setting at high tool revolving speed (2500 rpm) due to substantial reduction in axial thrust with higher stirring torque. The local deviation in thrust-torque during plunging was equally responsible with global mean thrust-torque in dwelling to dictate the joint strength and the weld ductility. The material stirring torque was better indicator of weld shear strength while pin assisted slot size highly correlated with tool thrust though tool plunging speed was highly crucial.

Variable pitch fan aerodynamic design for reverse thrust operation

Variable pitch low pressure ratio fans could enable higher propulsive efficiency and eliminate the need for heavy thrust reversers. In this paper, the effects of the fan rotor design on reverse thrust capability have been explored by varying key parameters of NASA’s Advanced Ducted Propulsor (ADP) while maintaining cruise point performance. Reverse thrust performance has been assessed using RANS single passage CFD of the variable pitch fan system with an extended domain. This computational approach has been validated using NASA Stage 51B, an experimental variable pitch fan test case. Predicted total pressure and total pressure ratios for this case were found to agree with test data within experimental error, except where large tip region separations occurred at high incidence. Applying individual variations to rotor pitch-to-chord, radial loading distribution, and chordwise camber distribution generated changes to rotor incidence, blockage, and peak surface Mach numbers in reverse operation. An increase in gross reverse thrust of up to 8% was achieved through reductions in rotor pitch-to-chord due to improved loading and lower shock Mach numbers. Controlling section camber distributions was used to move the rotor shock downstream and was found to increase reverse thrust by up to 3%. Increasing rotor loading at the mid span relative to the tip resulted in high tip incidence and a 50% reduction in gross reverse thrust across all rotor speeds. This suggests that tip loaded designs are required for high levels of reverse thrust.