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Low-speed characteristics of variable-sweep supersonic transport model with blended engine fuselage and engine-mounted tails
The static lateral- and directional-stability characteristics of a high-speed fighter-type airplane, obtained from wind-tunnel tests of a model, are presented. The model consisted of a thin, unswept wing of aspect ratio 2.3 and taper ratio 0.385, a body, and a horizontal tail mounted in a high position on a vertical tail. Rolling-moment, yawing moment, and cross-wind-force coefficients are presented for a range of sideslip angles of -5 deg. to +5 deg, for Mach numbers of 0.90, 1.45, and 1.90. Data are presented which show the effects on the lateral and directional stability of: (1) component parts of the complete model, (2) modification of the empennage so as to provide different heights of the horizontal tail above the wing plane, (3) angle of attack, and (4) dihedral of the wing.
This paper describes the turbofan engine architecture for SUbsonic Single Aft eNgine (SUSAN) Electrofan is, a transformative concept hybrid electric aircraft. SUSAN has a single tail-mounted turbofan engine that produces a portion of the requires thrust and drives a 20MW electric generator which in turn provides the power to the 16 electric propulsors located on the wing responsible for producing the remainder of the required thrust. The atypical operation of the turbofan due to the large levels of power extraction from the low pressure turbine (LPT) is described here. This paper investigates the most efficient engine architecture to enable this unique operation, as well as exploring natural gas as an alternative fuel for SUSAN.
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Experimental results showing the static longitudinal-stability and control characteristics of a model of a fighter airplane employing a low-aspect-ratio unswept wing and an all-movable horizontal tail are presented. The investigation was made over a Mach number range from 0.60 to 0.90 and from 1.35 to 1.90 at a constant Reynolds number of 2.40 million, based on the wing mean aerodynamic chord. Because of the location of the horizontal tail at the tip of the vertical tail, interference was noted between the vertical tail and the horizontal tail and between the wing and the horizontal tail. This interference produced a positive pitching-moment coefficient at zero lift throughout the Mach number range of the tests, reduced the change in stability with increasing lift coefficient of the wing at moderate lift coefficients in the subsonic speed range, and reduced the stability at low lift coefficients at high supersonic speeds. The lift and pitching-moment effectiveness of the all movable tail was unaffected by the interference effects and was constant throughout the lift-coefficient range of the tests at each Mach number except 1.90.
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Explore the source record for details and available documents.
Explore the source record for details and available documents.
As extensive experimental campaign was performed to study the aeroacoustic installation effects of an open rotor with respect to both a conventional tube and wing type airframe and an unconventional hybrid wing body airframe. The open rotor rig had two counter rotating rows of blades each with eight blades of a design originally flight tested in the 1980s. The aeroacoustic installation effects measured in an aeroacoustic wind tunnel included those from flow effects due to inflow distortion or wake interaction and acoustic propagation effects such as shielding and reflection. The objective of the test campaign was to quantify the installation effects for a wide range of parameters and configurations derived from the two airframe types. For the conventional airframe, the open rotor was positioned in increments in front of and then over the main wing and then in positions representative of tail mounted aircraft with a conventional tail, a T-tail and a U-tail. The interaction of the wake of the open rotor as well as acoustic scattering results in an increase of about 10 dB when the rotor is positioned in front of the main wing. When positioned over the main wing a substantial amount of noise reduction is obtained and this is also observed for tail-mounted installations with a large U-tail. For the hybrid wing body airframe, the open rotor was positioned over the airframe along the centerline as well as off-center representing a twin engine location. A primary result was the documentation of the noise reduction from shielding as a function of the location of the open rotor upstream of the trailing edge of the hybrid wing body. The effects from vertical surfaces and elevon deflection were also measured. Acoustic lining was specially designed and inserted flush with the elevon and airframe surface, the result was an additional reduction in open rotor noise propagating to the far field microphones. Even with the older blade design used, the experiment provided quantification of the aeroacoustic installation effects for a wide range of open rotor and airframe configurations and can be used with data processing methods to evaluate the aeroacoustic installation effects for open rotors with modern blade designs.
High lift characteristics of variable-sweep transport model with blended engine-fuselage and engine-mounted horizontal tails
A limited study has been conducted to establish the performance and noise characteristics of a low design tip speed (168 m/s, 550 ft/sec) low pressure ratio (1.04) variable pitch fan which was tested in the Langley 30 X 60 tunnel. This fan was designed for minimum noise when installed in the tail mount location of a twin engine aircraft which normally has both nose and tail mounted propulsors. Measurements showed the fan noise to be very close to predictions made during the design of the fan and extremely low in level (65 dBA at 1000 ft) with no acoustic treatment. This is about 8 dB lower than the unshrouded 2 blade propeller normally used in this installation. On the basis of tests conducted during this program, it appears that this level could be further reduced by 2 dBA if optimized acoustic treatments were installed in the fan duct. Even the best of the shrouded propellers tested previously were 7 dB higher in level than the Q-Fan without acoustic treatment. It was found that the cruise performance of this fan was within 5% of the predicted efficiency of 72%. Evaluation of the performance data indicated that disturbances in the inflow to the fan were the probable cause of the reduced performance.
The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached at Mach numbers above 0.90. Longitudinal dynamic stability calculations show that the longitudinal control is good at angles of attack below the unstable break in the static pitching-moment curves, but a typical corrective control applied after the occurrence of neutral stability has little effect in averting pitch-up.
The longitudinal aerodynamic characteristics of a full scale, single engine general aviation airplane equipped with an angle of attack limiting concept which used a tail mounted spoiler for stall prevention were investigated. The tail spoiler was deployed as a nonlinear function of changes in angle of attack in a narrow angle of attack range immediately preceding the stall angle. Spoiler deployment produced a nose down moment which limited the nose up trim capability of the elevator to an angle of attack below the stall angle of the airplane. The tail spoiler system produced significant increases in longitudinal stability and elevator hinge moments near the stall. The effects of power, flaps, and center of gravity position are also studied.
An experimental investigation has been conducted to examine the mechanisms of vortex-fin interaction on a twin-fin configuration. The investigation included a parametric study of the effect of tail location. The vortices were generated by a 76 deg sharp-edged delta wing with vertical tails mounted behind the wing. The model included both a dynamically-scaled flexible tail and a pressure instrumented rigid tail. Surface oil-flow patterns, off-body laser light sheet visualizations, aerodynamic load measurements, mean and unsteady flexible tail response, and unsteady tail surface pressure measurements were obtained. The results show that the tail location did not affect the upstream trajectory of the delta wing vortex. The tail location did affect the location of vortex breakdown, the global structure of the flow field, the aerodynamic loads, and the fin buffeting levels. The buffeting levels were reduced as the fins were moved laterally toward the vortex core trajectory. Two distinct peaks were observed in the pressure excitation spectra in the post-breakdown flow. Finally, the presence of the flexible tail opposite the rigid pressure tail altered the pressure measurements at one angle of attack.
A wind-tunnel investigation has been conducted to determine the effects of an unconventional tail arrangement on the subsonic static longitudinal and lateral stability characteristics of a model having a 63 deg sweptback wing of aspect ratio 3.5 and a fuselage. Tail booms, extending rearward from approximately the midsemispan of each wing panel, supported independent tail assemblies well outboard of the usual position at the rear of the fuselage. The horizontal-tail surfaces had the leading edge swept back 45 deg and an aspect ratio of 2.4. The vertical tail surfaces were geometrically similar to one panel of the horizontal tail. For comparative purposes, the wing-body combination was also tested with conventional fuselage-mounted tail surfaces. The wind-tunnel tests were conducted at Mach numbers from 0.25 to 0.95 with a Reynolds number of 2,000,000, at a Mach number of 0.46 with a Reynolds number of 3,500,000, and at a Mach number of 0.20 with a Reynolds number of 7,000,000. The results of the investigation indicate that longitudinal stability existed to considerably higher lift coefficients for the outboard tail configuration than for the configuration with conventional tail. Wing fences were necessary with both configurations for the elimination of sudden changes in longitudinal stability at lift coefficients between 0.3 and 0.5. Sideslip angles up to 15 deg had only small effects upon the pitching-moment characteristics of the outboard tail configuration. There was an increase in the directional stability for the outboard tail configuration at the higher angles of attack as opposed to a decrease for the conventional tail configuration at most of the Mach numbers and Reynolds numbers of this investigation. The dihedral effect increased rapidly with increasing angle of attack for both the outboard and the conventional tail configurations but the increase was greater for the outboard tail configuration. The data indicate that the outboard tail is an effective roll control.
Conceptual designs of the propulsor modules for the SUSAN electro-fan aircraft are sought after. Shaft power requirement is reduced by using boundary layer ingesting propulsion technology. There are several obstacles to designing feasible BLI propulsion systems such as inlet distortion, flow blockage from ingesting low-momentum flow, strong coupling between airframe and propulsion systems. Thus, the high fidelity CFD tool is indispensable to assess the performance of the propulsion systems and evaluate the inlet profiles during the design process. Consequently, the CFD data is used to update the inputs for the NPSS model from the initial stage of the system design. Hence, the inlet sizing, fan design, and estimation of the power saving are carried out. Various installation concepts of the mail-slot nacelle, such as under-/over-wing and trailing edge configurations, are investigated for the wing-mounted turbo-electric distributed propulsor module, and their power-saving is evaluated. As for the tail-mounted turbofan engine, the fan diameter and inlet captured area is determined based on the CFD profiles. An appropriate area ratio of the bypass and core ducts is derived from meeting the target bypass ratio from the system design. The baseline fan stage is analyzed by full annulus URANS CFD to assess the efficiency penalty due to the ingested pressure and swirl distortions.
Conceptual designs of the propulsor modules for the SUSAN electro-fan aircraft are sought after. Shaft power requirement is reduced by using boundary layer ingesting propulsion technology. There are several obstacles to designing feasible BLI propulsion systems such as inlet distortion, flow blockage from ingesting low-momentum flow, strong coupling between airframe and propulsion systems. Thus, the high fidelity CFD tool is indispensable to assess the performance of the propulsion systems and evaluate the inlet profiles during the design process. Consequently, the CFD data is used to update the inputs for the NPSS model from the initial stage of the system design. Hence, the inlet sizing, fan design, and estimation of the power saving are carried out. Various installation concepts of the mail-slot nacelle, such as under-/over-wing and trailing edge configurations, are investigated for the wing-mounted turbo-electric distributed propulsor module, and their power-saving is evaluated. As for the tail-mounted turbofan engine, the fan diameter and inlet captured area is determined based on the CFD profiles. An appropriate area ratio of the bypass and core ducts is derived from meeting the target bypass ratio from the system design. The baseline fan stage is analyzed by full annulus URANS CFD to assess the efficiency penalty due to the ingested pressure and swirl distortions.
A joint U.S. Army/NASA flight investigation was conducted utilizing a single-rotor helicopter to determine the effectiveness of horizontally mounted tail boom strakes on directional controllability and tail rotor power during low-speed, crosswind operating conditions. Three configurations were investigated: (1) baseline (strakes off), (2) single strake (strake at upper shoulder on port side of boom), and (3) double strake (upper strake plus a lower strake on same side of boom). The strakes were employed as a means to separate airflow over the tail boom and change fuselage yawing moments in a direction to improve the yaw control margin and reduce tail rotor power. Crosswind data were obtained in 5-knot increments of airspeed from 0 to 35 knots and in 30 deg increments of wind azimuth from 0 deg to 330 deg. At the most critical wind azimuth and airspeed in terms of tail rotor power, the strakes improved the pedal margin by 6 percent of total travel and reduced tail rotor power required by 17 percent. The increase in yaw control and reduction in tail rotor power offered by the strakes can expand the helicopter operating envelope in terms of gross weight and altitude capability. The strakes did not affect the flying qualities of the vehicle at airspeeds between 35 and 100 knots.