NASA Technical Memorandum 81298
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IWVSA National Aeronautics and Space Administration September I98I NASA Technical Memorandum 81298
Quiet Short-Haul Research Aircraft Familiarization Document, Revision 1
Joseph C. Eppel, Ames Research Center, Moffett Field, California
NASA National Aeronautics and Space Administration Ames Research Center Moffett Field California 94035 TABLE OF CONTENTS
NOMENCLATURE
1. INTRODUCTION
2. AIRCRAFT DESCRIPTION.
3. AIRCRAFT SPECIFICATIONS 3 3.1 Weights and Inertias 3 3.2 Center of Gravity 4 3.3 Dimensions and General Data 4 3.4 Control Surface Areas, Deflections and Rates 5 3.5 Cargo Area 5 3.6 Landing Gear 5
4. STRUCTURAL DESIGN CRITERIA 6 4.1 Design Speeds 6 4.2 Limit Load Factors 6 4.3 Structural Load Factors 6 4.4 Service Life '. . . . 7
5. PERFORMANCE AND NOISE 7 5.1 Takeoff 7 5.2 Stopping Distance 7 5.3 Climb and Descent 7 5.4 Range and Endurance 7 5.5 Noise 8
6. PROPULSION 8 6.1 Configuration and Powerplant ~ 8 6.2 Powerplant 9 6.3 Nacelle and Powerplant Installation 9 6.4 Engine Performance .- 9 6.5 Acoustic Treatment 9 6.6 Nacelle and Engine Instrumentation 9
7. BOUNDARY-LAYER CONTROL SYSTEM 10
8. FLIGHT-CONTROL SYSTEM 10 8.1 Basic Flight-Control System 10 8.2 Stability Augmentation System 12 8.3 Electric Command System 12
9. FLIGHT-TEST INSTRUMENTATION SYSTEM 13 9.1 Sensors and Transducers 13 9.2 Tape Recorder 13 9.3 Time-Code Generator 14 9.4 Active Network Panels 14 9.5 Parameter List 14
iii 10. OPERATIONAL SUBSYSTEMS 14 10.1 Fuel System 14 10.2 Hydraulic System 15 10.3 Electrical System 15 10.4 Environmental Control System 15 10.5 Communications and Navigation Systems 15 10.6 Crew Station 16 10.7 Emergency Egress ' . . 17
11. GROUND SERVICING AND HANDLING 17 11.1 Ground Service Equipment 17 11.2 Fluid Requirements 17 11.3 Airplane Grounding 18 11.4 Servicing 18
12. DOCUMENTATION LIST 18
APPENDIX - LIST OF QSRA REPORTS 19
TABLES 21
FIGURES 26
IV NOMENCLATURE
ADF automatic direction finder / AEO all engines operating
BLC boundary layer control
BPR bypass ratio
CL lift coefficient
DLC direct lift control
6w,6t wheel deflection, throttle deflection
EPNdb equivalent perceived noise in dB
Fs stick force
IFF identification, friend or foe
IRIG interrange instrumentation group
KEAS knots equivalent air speed
MGT measured gas temperature
NL,NH fan rpm, core rpm
OEW operational empty weight (excludes data system, pilots)
SAS stability augmentation system
SFC specific fuel consumption
TMAX maximum thrust
USB upper surface blown
U^ gust loading, ft/sec e VQL climb velocity
VFLDG landing velocity
VHF very high frequency
VOR/ILS VHF omnirange, instrument landing system
v ZFW zero fuel weight (includes pilots)
<(>,6,qc roll angle, pitch angle, dynamic pressure
vi QUIET SHORT-HAUL RESEARCH AIRCRAFT FAMILIARIZATION DOCUMENT, REVISION 1
Joseph C. Eppel
Ames Research Center
1. INTRODUCTION
This document summarizes the Quiet Short-Haul Research Aircraft's (QSRA) design features and performance characteristics and describes its flight-test data-acquisition system. The information contained herein was. obtained during the design, ground test, and first 30 hr of flight testing, and should there- fore be treated as preliminary data. More recent data have been presented in reports listed in the appendix. Additional data will be reported as analyses progress.
Propulsive-lift technology is needed by government and industry to develop options for future U.S. civil short-haul transportation, to compete in foreign markets, and to improve military tactical airlift capability. The objective of NASA's Quiet Propulsive-Lift Technology (QPLT) program is to pro- vide data which: (1) will significantly reduce the technical risk associated with the development by industry of civil and military propulsive-lift trans- ports, and (2) will permit government regulatory agencies to establish real- istic criteria for certification of commercial propulsive-lift aircraft.
The QSRA (fig. 1) is the major element of the QPLT program. The airplane is a flight-research facility which is being used for advanced flight experi- ments. It has very high performance in the low-speed regime, very low noise levels, and a versatile flight-control system. The research application of the aircraft is focused on takeoff and landing and other terminal area opera- tions associated with the propulsive-lift mode of flight.
The QSRA was designed and built by the Boeing Commercial Airplane Company. It is a modified de Havilland C-8A Buffalo, with jaew wing and nacelles and four AVCO-Lycoming YF-102 engines. It employs the upper-surface-blowing propulsive- lift concept. Engine bleed air provides trailing-edge (aileron) blowing for boundary-layer control. The first flight occurred in July, 1978; the aircraft was delivered to Ames Research Center in August.
Cost was second only to safety in QSRA program priorities. Consequently, the Boeing contract contained performance goals rather than specifications. This permitted design tradeoffs between research capability and cost, with the intent to maximize the overall capability of the airplane while remaining within the project budget. The contractor was encouraged to suggest changes that resulted in increased research capability at equal cost or reductions in equipment salvaged from other airplanes (fuselage, engines, accessories, etc.) also reduced cost. This resulted in a very austere airplane, but one with a very high research capability. The Quiet Short-Haul Aircraft Office at Ames is responsible for manage- ment of the QSRA flight experiments program. Experiments are planned in various technical disciplines: aerodynamics, propulsion, stability and con- trol, handling qualities, avionics and flight-control systems, acoustics, loads and structures, operating systems, human factors, trailing-vortex phe- nomena, and airworthiness/certification criteria. Investigators from various organizations at Ames and other NASA centers participate in the definition and conduct of these experiments. However, it is NASA's intent that the QSRA be a national facility. Investigators from industry, the academic community, and other government agencies are encouraged to propose and participate in flight experiments. Two examples of outside experiments are:
1. A joint NASA/Navy flight research program was conducted using the QSRA to investigate the application of advanced propulsive-lift technology in the naval aircraft carrier environment. Unarrested landings and free deck takeoffs were successfully accomplished.
2. The Boeing Commercial Airplane Company is sponsoring an experiment which involves the testing of inboard spoilers being driven by samarian cobalt motors instead of the existing hydraulic actuators. This experiment will remain on the QSRA in parallel with the ongoing NASA flight research program.
2. AIRCRAFT DESCRIPTION
The Quiet Short-Haul Research Aircraft (fig. 2) is designed as a facility for flight research with very high levels of low-speed performance and low community noise levels. As the emphasis is on terminal area operation, the airplane is fixed in the low-speed configuration, with the leading-edge devices deployed and landing gear locked down. Virtually no attempt has been made to reduce drag for cruise performance, although to make the airplane more representative of a commercial transport, it incorporates such fuel- conservative features as a high-aspect ratio, moderately swept wing, a super- critical airfoil, and a high-performance inlet without acoustic rings. High- speed wind tunnel tests are being planned for FY .82-83 to study the problems associated with efficient cruise performance of QSRA-type propulsive-lift configurations. Advanced avionics equipment will be installed during the third quarter of 1981. This equipment is an integrated avionics system which provides a capability for terminal-area navigation, guidance, and control research on STOL aircraft.
The QSRA uses a de Havilland C-8A Buffalo fuselage and landing gear with modifications and a C-8A empennage, modified for powered elevator operation and fitted with SAS actuators. The primary feature is the new wing with internal fuel tanks, fitted with four new nacelles forward and over the wing which house Lycoming YF-102 turbofan engines. The new configuration details include the advanced airfoil, upper-surface-blowing powered lift, aileron boundary-layer control blowing, and an advanced flight-control system featur- ing electrically commanded fly-by-wire spoilers and flaps. Three-axis sta- bility augmentation is provided. An airborne flight-test instrumentation system is installed in the aft cabin. Due to its research nature, the air- plane carries only a pilot and copilot. There is provision for additional experimental payload in the cabin area. While space is abundant, the airplane is not configured to carry observers, only the two research pilots.
3. AIRCRAFT SPECIFICATIONS
The data presented herein represent the current configuration. However, ongoing changes have increased the OEW and ZFW. Also the pending installa- tion of advanced avionics equipment will affect the group weights.
3.1 Weights and Inertias 3.1.1 Airplane weights (July 1981)-
Takeoff gross weight, maximum 60000 Design landing weight at maximum 48000 (see table 1, p. 21) sink rate Zero fuel weight (ZFW) 40258 Operational empty weight (OEW) 37600 (No flight-test instrumentation) Maximum taxi weight 60000
3.1.2 Group weights - Ib X-C.G. - in. Wing 8742 389.2 Horizontal tail 614 906.9 Vertical tail 494 840.8 Body 6121 410.0 Main landing gear 2098 392.3 Nose landing gear 288 55.7 Nacelle and strut 4454 290.1 Total structure (22811) (396.1) Engines 4960 281.6 Engine accessories 317 255.8 Engine controls 92 221.7 Starting system 152 268.9 Total propulsion group (5791) (280.9) Instruments 198 86.3 Surface controls 1356 437.4 Hydraulics 958 388.0 Pneumatics 392 285.2 Electrical 1311 230.0 Electronics 491 173.8 Boundary-layer control 948 199.5 Furnishings 331 354.7 Cargo handling 101 331.8 Emergency equipment 308 290.5 Air conditioning 128 142.5 Anti-icing 20 545.5 Total fixed equipment (6562) (310.8) Ballast 872 837.0 Exterior paint 115 439.0 Calculated to actual increment -321 107.2 Manufacturer empty weight 35830 Standard and operational items 770 220.4 Operational empty weight 36600 372.1 Flight-test equipment 2120 435.5 Maximum zero fuel weight 38720 375.5
3.1.S Airplane inertias (maximum gross weight)-
I (roll) 193,660 slug ft2 XX I (pitch) 272,620 slug ft2
2 I..tj ,£» (yaw) 420,700 slug ft
2 Izz (product) 23,210 slug ft The variation in inertia with airplane gross weight is shown in figures 3(a), (b), (c), and (d).
3.2 Center of Gravity
The weight, balance, and main gear limit for the QSRA are shown in figure 4.
3.3 Dimensions and Genera"! Data
Horizontal Vertical Wlflg tail tail
Area, ft2 600 233 152 Span, ft 73.5 32.0 13.58 Chord - root, in. 150.7 100 168 Chord-tip, in. 45.2 75 100 Mean aerodynamic chord, in. 107.4 88 137 Aspect ratio 9.0 4.4 1.22 Incidence, deg 4.5 3° L.E. UP - Dihedral, deg 0 0 - Sweep C/4, deg 15.0 3.0 17.37 Taper ratio 0.30 0.75 0.60 Thickness ratio - side of 18.54 14 14 body, % Thickness ratio - tip, % 15.12 12 14
3.4 Control Surface Areas, Deflections and Rates
Area Deflection, Maximum rate, ft2/APL deg deg/sec
Aileron 32.2 -20, +50 ±40 USB flaps 105 0, 66 0-30, ±10 30-66, ±6.74 Outboard flaps 40.2 0, 59 ±5 Spoilers 33.7 60 up ±50 Direct lift control - biased to 13° up with ±13° travel about that point Elevator 81.6 -25, +15 ±45 Rudder 60.8 ±25 ±35
3.5 Cargo Area
The main cabin area of the QSRA, shown in figure 5, is extensive, but due to existing installations, the clear floor area is limited to approximately 7 ft by 10 ft. Restrictions in access limit package sizes to 30 in. wide and 48 in. high.
3.6 Landing Gear
Main landing gear-
Wheels per side 2 Wheel spacing 25 in. Tire size, type, rating 37 x 14, 15 diam, 22 PR Inflation pressure 90 psi Strut deflection 24 in. Nose landing ^ear- Wheels 2 Wheel spacing 18 in. Tire size, type, rating 8.90 x 12.50, Type 3, 6-ply Inflation pressure 42 psi Strut deflection 18 in. 4. STRUCTURAL DESIGN CRITERIA
4.1 Design Speeds
The design £tructural limit speeds are listed below:
V Maximum ground operation speed 120 knots G
VD Design dive speed 190
VMQ, VG Maximum operating and cruise speed 160
V. Design maneuvering speed 142 A
VD Gust slowdown speed 137 D
VMr Minimum controllable airspeed 55 (1 g flight CEI)
Von Speedbrake extend limit 190 3D VT, Flap extend speeds are listed in the following table: r
_,. , . _. Outboard USB Limit air- Flight configuratioc n ,., .. ,_, , , , & flaps, deg flaps, deg speed, knots
Go-around from CTOL 30 0 140 Takeoff, CTOL landing 59 0 130 Go-around from STOL 59 <40 120 STOL landing 59 >40 100
4.2 Limit Load Factors
The QSRA is limited to the maneuver load factors in the following table. V-n diagrams are provided in figure 6.
„ . ... Positive load Negative load Configuratio f n , 7 factor, g factor, g
Flaps up, to 160 knots 2.25 0.5 Flaps up, above 160 knots 2.25 0.0 Flaps extended 2.0 0.0
4.3 Structural Load Factors
The structural design loads are 1.74 (1.5 x 1.15 = 1.73) times the limit loads. The 1.15 factor was used in lieu of a structural test program. 4.4 Service Life
The QSRA has been designed for a service life of not less than 2000 flights or 500 flight hours without replacement or overhaul of major components.
5. PERFORMANCE AND NOISE
5.1 Takeoff
Takeoff performance for QSRA is shown in figure 7. The data provided for field length are based on the longest of:
1. 1.15 times the distance to clear a 35-ft obstacle with all engines operating.
2. The distance to 35-ft altitude with an outboard engine failed at the critical failure speed.
3. The distance to accelerate to the critical speed and stop. Airplane configuration is with takeoff flaps (outboard flaps 59°, USB flaps retracted), and assumes thrust is not increased on remaining engines after an engine failure.
5.2 Stopping Distance
Landing ground distance is given in figures 8(a) and 8(b) for various approach speeds, flightpath angles, and temperatures. Distances have been multiplied by 1.67 as is done for Federal Air Regulation distances. Nonfac- tored distances are provided in figure 8(c).
5.3 Climb and Descent
5.3.1. Climb- Figure 9(a) gives fuel, distance, and time to climb for cruise configuration, standard day, with Boundary-Layer Control (BLC) switches off. Figures 9(b), (c), and (d) give climb performance at sea level, 5,000 ft, and 10,000 ft with BLC switches off. Figure 9(e) gives climb performance at 15,000 ft with the BLC switches off.
5.3.2 Descent- The descent capability of the QSRA in the cruise configu- ration at 50% NL is shown in figure 10.
5.4 Range and Endurance
The range for a ferry mission for a takeoff gross weight of 48,820 Ib (10,100 Ib of fuel) is 256 n. mi., with an elapsed time of 128 min. This mission is accomplished as shown below: Segment Fuel used» Time' Distance, lb min n. mi. Taxi 330 15.0 Takeoff, climb to 1500 ft 190 .9 Climb to 15,000 ft 1030 8.6 19 Cruise 6220 8.2 220 Descent to 1500 ft 230 7.2 17 Land 300 3.8 Taxi 22Qa 10.0 Totals 8520 127.5 256 Taken from reserve.
A 30-min reserve (1,800 lb of fuel) was used. BLC switches are turned off except for landing. A schematic of a STOL mission is shown in figure 11. The airplane weight is 45,000 lb for this case. There is enough fuel (1,800 lb reserve) for 12 cycles.
5.5 Noise
The community noise generated by QSRA estimated at first flight were as follows: 500-ft sideline noise
Takeoff, 50,000-lb aircraft, 2,000-ft runway 91 EPNdB Landing, 48,000-lb aircraft, y = -7.5°, 70 knots 89 EPNdB
The area within the 90 EPNdB footprint for a combined takeoff and landing, scaled to a 150,000-lb airplane, is estimated to be approximately 1 mi2.
6. PROPULSION
6.1 Configuration and Powerplant
The QSRA propulsion system consists of four AVCO-Lycoming YF-102 engines mounted in nacelles above and forward of the wings as shown in figure 12. These prototype engines, obtained from the Air Force A-9A program, were refur- bished and updated by the engine manufacturer. The principal elements of this update consisted of a fan containment ring, combustor case bleed ports, and new oil coolers. 6.2 Powerplant
A cutaway view of the engine is shown in figure 13(a). The engine is basically a T-55 gas producer with the addition of a two-stage turbine driving a single-stage fan through 2.3 to 1 reduction gears. Engine geometry, dimen- sions, and uninstalled performance are shown in figure 13(b). The engines are provided with an overspeed protection system that senses the power turbine shaft speed and shuts off the fuel at the fuel control if the turbine exceeds 103% of its design speed.
6.3 Nacelle and Powerplant Installation
The layout of the nacelle is shown in figure 14(a). The QSRA nacelle and powerplant installation are unusual in that the engine serves as an integrated part of the nacelle structure, supporting the inlet and nose cowl, core cowl, primary nozzle, and engine-driven accessories (fig. 14(b)). The remainder of the nacelle is the structural cowl and nozzle assembly, which mounts on the wing and supports the engine buildup. The nozzle assembly mixes the fan and core air and terminates in a D-shaped opening that shapes the flow into a thin sheet, which in turn flows over the wing and upper-surface-blowing (USB) flaps to provide lift.
6.4 Engine Performance
The relationship of thrust and fan speed for the YF-102 engines is shown in figure 15(a), which also shows the relationship between core speed and fan speed. The effect of ambient temperature on installed thrust is shown in figure 15(b).
6.5 Acoustic Treatment
The location of the acoustic liners in the nacelle is shown in figure 16. The liners in the fan duct are found on the structural cowl and the core cowl, and are composed of perforated face sheets, aluminum honeycomb cores, and solid back sheets. They cover about 30 in. of the duct length and are esti- mated to provide 12 PNdB of aft fan attenuation. The inlet lining is of double wall construction with perforated aluminum face sheet and septum, aluminum honeycomb cores, and solid aluminum backing sheets.
6.6 Nacelle and Engine Instrumentation
Each inlet has four static pressure taps located 90° apart slightly down- stream of the throat area. The liners are not removable at this time and therefore the ability to add inlet instrumentation is limited. Serial number 2 engine is installed on the left inboard location and contains the instrumentation described in section 9. The operation of this engine was docu- mented during the ground test and correlated with each of the other engines and with earlier test cell data. 7. BOUNDARY-LAYER CONTROL SYSTEM
The Boundary-Layer Control (BLC) system improves the performance of the QSRA by enabling operation with good handling qualities and increased angle- of-attack margins down to very low airspeeds. Each outboard engine provides air for an opposite aileron, as shown in figure 17.
Two key elements of the BLC system are the mixing ejector and its asso- ciated servo-pressure regulator valve, which are located in the outboard nacelles behind the fan and below the core. The ejector, shown in figure 18, mixes high-pressure core bleed air and fan air to provide constant duct pres- sure with varying engine speeds. The ejector itself uses a fixed geometry mixing section with an elliptical centerbody, and a series of high-pressure nozzles located circumferentially between the duct wall and the centerbody. The core bleed is limited to 10% by the high-pressure nozzles, while the fan bleed is limited to 3% by the duct size. Figure 18 shows the net blowing momentum of a BLC nozzle as provided by the ejector on the upper curve without regulation. The curve on the right shows the maximum available engine thrust (temperature limit) for various values of core bleed airflow. The lower dashed line shows the value with the pressure regulator valve in operation, sensing pressure downstream of the ejector and controlling the high-pressure bleed air to the ejector. The regulator gives no bleed at high engine speeds where the fan flow is adequate, and up to 10% bleed at low speeds. The ensuing loss of thrust due to core bleed only occurs at low throttle settings where high power is not required.
The regulated air is routed through ducts to a series of adjacent nozzles which direct the flow over the ailerons, significantly improving the control effectiveness during low-speed, high-lift operation. Automatic cutoff of the regulated core air is provided to protect the engines from overbleeding at high power, which could occur with a blown BLC duct.
8. FLIGHT-CONTROL SYSTEM
8.1 Basic Flight-Control System
The Quiet Short-Haul Research Aircraft flight-control system is shown schematically in figure 19(a), and pictorially in figure 19(b). The airplane can be flown from either the right- or left-hand seat, although the cockpit instrument panel and controls are optimized for flying from the left seat. In addition, the control wheels are fitted with force and position sensors, which furnish inputs to the stability and control augmentation system. All primary flight controls are conventional. Power levers are located in an overhead throttle console.
8.1.1 Lateral control system- The lateral control system is schematically shown in figure 20(a). Roll control is achieved using ailerons and spoilers. The centering detent and wheel force gradient are provided mechanically. Aileron trim is provided by electrically repositioning the feel detent. The
10 ailerons are drooped as a function of outboard flap position and operate about that point. Deflections of the ailerons and spoilers as a function of wheel position are shown in figure 20(b). Roll acceleration for full wheel deflec- tion as a function of airspeed is shown in figure 20(c).
8.1.2 Longitudinal control system- The longitudinal control system is shown schematically in figures 21(a) and 21(b). Longitudinal trim is provided by moving the neutral position of the feel and centering unit. Manual rever- sion trim is provided by a manually-driven trim tab. A geared balance tab is utilized to reduce pilot loads during manual reversion operation. Predicted pitch response to full elevator deflection vs airspeed is shown in figure 21(c).
8.1.3 Directional control system- The rudder control system is shown schematically in figure 22(a). The double-slotted rudder has a SAS actuator for automatic flight-control inputs. The trim function is implemented through the feel unit neutral position. The predicted yaw acceleration with maximum rudder deflection vs airspeed is shown in figure 22(b).
8.1.4 Flap system- The outboard flaps are conventional double-slotted flaps and primarily develop high lift. They operate through an electric com- mand system from the outboard flap lever in the overhead console. In addition to symmetric operation, the outboard flaps can provide a lateral trim function for engine-out operation by retracting the outboard flap on the side opposite the failed engine. This function is controlled by the pilot through a switch on the windshield center post.
The outboard flap electric command computers also detect asymmetry, and lock the flaps in their present position if the asymmetry exceeds 10° for more than 1 sec, unless that asymmetry is commanded through the lateral trim system.
The trailing-edge flap system is composed of the inboard and center upper-surface-blown (USB) flaps and the outboard flaps, and is shown sche- matically in figure 23. The USB flaps are electrically commanded, and operate in unison to provide lift and thrust vector control. Deflections from 0° to 30° are controlled by the position of the USB flap handle on the overhead con- sole, while deflections from 30° to 66° are selected using a thumb switch on the number one and number four throttle levers. The USB flap computers lock the flaps in their present position if an asymmetry of more than 10° exists for more than 0.5 sec.
8.1.5 Spoiler system- The spoiler system uses inputs from the speedbrake handle, the pilot's wheel, and the longitudinal stability augmentation system (SAS) computer, which produces a spoiler position as a function of throttle position for direct lift control. These signals are operated on by an electric command computer, which drives the four spoiler panels (two on each wing for- ward of the outboard flaps) using electro-hydraulic actuators on each panel. It is planned as an experiment to use electro-mechanical actuators on the inboard panels. The speedbrake function produces symmetric spoiler deployment from 0° to 60° proportional to the speedbrake handle. The roll control func- tion operates by driving the spoilers on the appropriate wing upward to pro- duce a rolling moment to augment the ailerons. Direct Lift Control (DLC), when selected, positions the spoilers at 13°. When the throttles are advanced, the
11 spoilers drop from that point; when the throttles are retarded, the spoilers move upward. These movements produce a rapid change in the lift to quicken the flightpath response to throttle changes. The deflections are washed out to the bias position after the initial response. This option is selectable only when the USB flaps are at 30° or more (landing configuration). Computation for mechanization of the DLC function is performed in the longitudinal SAS computer.
8.2 Stability Augmentation System (SAS)
The QSRA is provided with three-axis stability augmentation with each axis being independent. Principal inputs and outputs are shown in figure 19(a).
The roll and yaw stability augmentation systems provide improved airplane handling qualities resulting in a decrease in pilot workload. Roll-mode and spiral-mode augmentation is provided in the lateral axis, with dutch roll damping and turn coordination provided in the directional axis. In addition, control wheel position information is used in the lateral axis to improve the linearity of the airplane roll response to wheel inputs. An option added to the lateral stability augmentation system incorporates, at the pilot's option, a roll rate command/attitude hold system. When the airplane is in flight, a roll rate is commanded proportional to the wheel position. When the wheel is in detent, the existing roll attitude is maintained. If the roll angle is less than 3°, the system will level the airplane and maintain zero roll angle.
The stability augmentation systems are nonredundant. Failure transients are kept to a minimum by rate and position limiting of the electro-hydraulic SAS servo actuators.
Pitch augmentation is provided in the form of a rate command/attitude hold system. The pitch augmentation system also provides the DLC function by processing throttle position signals and sending symmetric deflection commands to the spoiler control electronics. The pitch SAS computer is a digital mech- anization and drives the elevator through the elevator SAS servo.
8.3 Electric Command System
8.2.1 Spoiler system- The spoiler electric command system is composed of two isolated separate subsystems. The first of these drives the outboard spoilers on each wing; the other drives the inboard spoilers. The spoiler computers have dual signal paths and process the multiple command signals from the pilot's wheel, the spoiler lever, and the pitch SAS computer to drive the two servos. The components of the system are shown schematically in figure 24.
8.3.2 Upper-surface-blown flap system- The electric, USB flap command system consists of two separate and isolated subsystems, one driving the inboard flaps and the other, the center flaps. The computers incorporate single-channel signal processing and dual-channel drive electronics for the two electrohydraulic flap actuators associated with the subsystem. If the position feedback signals disagree by a predetermined amount, a monitor
12 circuit detects the asymmetry and locks those two servos in their current position. The components of the system are shown schematically in figure 25.
8.3.3 Outboard flap system- The outboard flap system consists of a single system driving the two outboard flaps. The computer mechanization pro- vides single-channel signal processing, from the outboard flap handle and the flap trim switch, with dual-channel servo-drive electronics for the two flaps. Monitoring of symmetry is mechanized as in the USB computers. Figure 26 shows the components of this system.
9. FLIGHT-TEST INSTRUMENTATION SYSTEM (FTIS)
The QSRA is equipped with a high-speed data system comprised of trans- ducers, signal conditioning equipment, a telemetry transmitter, and a tape recorder. This system measures, transmits, and records significant airplane parameters for real-time or after-the-fact analysis.
The FTIS, illustrated in block diagram form in figure 27, operates as follows. Data from the transducers are transmitted to the analog and digital network panels, which provide the necessary signal conditioning. The condi- tioned data then passes to the remote multiplexer/digitizer units (RMDU), which adjust the gains to a programmed level, provide analog to digital con- version, and encode the data in a pulse-code modulation, serial bit stream. Separate precision low-voltage power supplies located in the analog network panels furnish transducer excitation power where required. The FTIS contains a time-code generator which furnishes time correlation for the system, which can be synchronized with ground Interrange Instrumentation Group (IRIG) time. The data are recorded on a standard 14-track, airborne, magnetic tape recorder. An interface is available for in-flight transmission of data from one RMDU via L-band telemetry to a ground station for real-time data monitoring. A more detailed description of system elements follows.
9.1 Sensors and Transducers
Numerous sensors and transducers, such as thermocouples, pressure trans- ducers, strain gauges, servos, potentiometers, accelerometers, etc., are installed on the QSRA to measure parameters of interest.
9.2 Tape Recorder
The tape recorder is an Astro-Science (Bell and Howell) airborne wide- band FM recorder model MARS-144 (LT)-3D. The unit is a 14-track analog recorder, and uses a 10-in. reel of magnetic tape. A PCM data bit stream from each RMDU and the time-code generator output are recorded on separate tracks. The tape recorder is the limiting component in the FTIS due to its bit-packing density limitation as related to amount of tape available, speed of running, and length of record required for flight data recording during tests. The
13 recorder is operated at 7-1/2 in./sec, which allows about 2 hr of full-time data recording, which will cover one data flight.
9.3 Time-Code Generator
The time-code generator is a Datametrics Model SP-375 Airborne Synchro- nized Generator, which produces an IRIG B output for recording on the tape recorder. The unit will be synchronized with the ground time-code generator for flight tests and can be synchronized with radio station WWV.
9.4 Active Network Panels
The analog and digital active network panels process the transducer signals to make them compatible with the RMDU's. They also contain a Tecnetics power supply module, which provides ±5 V d.c. for transducer excitation.
9.5 Parameter List
The FTIS currently records the parameters listed in tables 2 and 3 (pp. 22-25). There is limited capability to add additional parameters if required, for specific experiments.
10. OPERATIONAL SUBSYSTEMS
10.1 Fuel System
The QSRA fuel system is diagrammed in figure 28. It is comprised of two fuel tanks contained in the wing from which four fuel boost pumps feed two fuel manifolds, one for each pair of left-hand and right-hand engines. The engines can also function on engine-driven fuel pumps using suction feed. Fuel levels between the left and right tanks can be balanced by shutting off the boost pumps on the side with more fuel remaining. Each tank holds 5,100 Ib of JP-5 or Jet-A fuel, and is vented to the atmosphere. Two auxiliary fuel tanks capable of holding an additional 5,400 Ib of fuel were installed in the cabin during the spring of 1979. As shown in figure 28, there is a single fueling inlet. During fueling the float-actuated valves are commanded open by an electrical switch and automatically closed when the tanks are full (and can- not be commanded open). During the fueling process, the fuel takes two paths, to the wing tanks and to the auxiliary tanks. The fuel transfer switch simul- taneously opens the wing floatation valves and starts the pump located on the auxiliary tank. Fuel can be transferred both on the ground and in the air.
The airplane can be fueled either by conventional pressure fueling or by over-the-wing gravity filler ports. For pressure fueling, automatic float- operated shut-off valves are provided. Defueling can be accomplished using the suction capability of ground fueling equipment.
14 10.2 Hydraulic System
The QSRA hydraulic system is shown schematically in figure 29. Hydraulic power is supplied by three separate and isolated hydraulic systems at 3000 psi. Systems A and B are each pressurized by two engine-driven pumps. The system A pumps are driven by engines 1 and 2, and the system B pumps are driven by engines 3 and 4. System C is pressurized by an a.c. motor-driven pump run full time. System fluid is MIL-H-5606.
The A and B systems are configured so that their power demands are approximately equal in size. One pump, driving system A or B, will meet all the system requirements of the given system should a single associated engine or pump failure occur. System C drives the aileron actuators only.
All primary control surface actuators are of a dual tandem configuration, and each actuator is supplied from two different power systems. Thus, normal surface actuation rates can be maintained in the event a single power system loss occurs. In the event of a double system failure, lateral control is maintained since the aileron actuators are powered from systems A, B, and C.
10.3 Electrical System
The QSRA electrical system is shown schematically in figure 30. Four 15 KVA, 115 V, 400 Hz generators are driven by the engines through constant speed drives. The two left generators are connected together to form one system; the two right generators provide a separate system. Automatic power transfer allows one generator to pick up the entire load for each system in the event of a generator or engine failure. Transformers on each generator provide 26 V a.c. power. Two transformer/rectifier units produce 28 V d.c. power. Manually-controlled bus-tie relays transfer all power between left and right 115 V a.c. systems if both generators in one system should fail.
10.4 Environmental Control System (ECS)
The environmental control system (ECS) for the QSRA is shown schematically in figure 31. This system provides heating, cooling, and ventilation to the flight deck and rear cabin. Engine bleed air, precooled by fan flow, or ground-start cart air is supplied to an air-conditioning pack which provides cooled air. This conditioned air and/or ram air from a scoop on the airplane nose is mixed with engine bleed air to provide the desired temperature in the crew areas. The percentage of air diverted to the aft cabin is controlled by a manual valve. The air to the flight deck is run through a muffler to reduce noise. The system is capable of holding temperatures in the cockpit to 80° F on a hot day.
10.5 Communications and Navigation Systems
The QSRA has one UHF and one VHP transceiver for communications. The following navigation and guidance systems are provided:
15 Automatic direction finder (1) Tacan (1) VOR receiver (2) Radar transponder (1) Marker beacon/glideslope receiver (1) J-2 compass (1)
10.6 Crew Station* 10.6.1 Cockpit arrangement- The QSRA has side-by-side seats for a pilot and copilot. The seats are not ejection seats and crew emergency egress is described in section 10.7. The airplane may be flown from either seat but operation of the longitudinal SAS and some displays make the left seat prefer- able. In addition to the instrument panel, there is a center console and an overhead console.
10.6.2 Instrument panel- The instrument panel is shown in figure 32(a), and contains the following:
Attitude indicators (2) Oil pressure and temperature Angle-of-attack indicators (2) indicators (4) Sideslip angle indicator Fuel flow indicators (4) Airspeed indicators (3) BLC pressure indicators (4) Course selectors (2) Flap position indicators Turn and slip indicators (2) Outboard Vertical speed indicators (2) USB center Barometric altimeters (2) USB inboard Normal acceleration indicator d.c. meters Tacan indicator a.c. meters Radar altimeters indicator Surface position indicator Battery temperature monitor (aileron, rudder, spoilers) Brake pressure indicator Surface position indicator Hydraulic pressure indicators (2) (elevator) Hydraulic quantity indicators (2) Annunciator Panel (60 positions) Fan speed indicators (4) Master caution lights (2) Core speed indicators (4) Master fire warning lights (2) Mixed gas temperature indicators (4) Miscellaneous controls, switches 10.6.3 Overhead console and eyebrow panel- The overhead console is shown in figures 32(a) and (b) and contains the following: Fire switches (4) Lighting control panel Flap levers (USB and outboard) (2) Hydraulic control panel Speed brake lever ECS panel Thrust levers Bleed air and ice-control panel Gust lock lever The eyebrow panel above the windshield contains:
Rudder trim switches (2) Elevator trim cutout switches (2) Rudder trim position indicator
*Crew station will be affected by advanced avionics installation.
16 10.6.4 Center console- The center console (aisle stand) is shown in figure 32(c) and contains the following control panels:
IFF Interphone (2) ADF Antiskid VHF communications Marker/beacon glideslope VHF (VOR/ILS) Flight-test instrumentation Tacan VHF communications Stability Augmentation System (SAS) Roll SAS wheel steering
10.7 Emergency Egress
The QSRA is provided with seven escape doors or hatches, the locations of which are shown in figure 33. The normal entrance/exit ladder under the aft fuselage can be used for bailout. The primary bailout exit is through the jettisonable hatch in the cockpit floor. Ground exit is possible through the pilot's and copilot's jettisonable side windows or the pilot's overhead escape hatch. Emergency exit hand slides are provided on the left and right sides of the cockpit. Side cabin escape hatches are also available.
11. GROUND SERVICING AND HANDLING
11.1 Ground Service Equipment
In order to operate the QSRA, the following facilities are needed:
1. Ground electrical cart (type MD-3A or equivalent) - This unit fur- nishes 115 V a.c. (45 kW) and 28 V d.c. (60 kW) for airplane ground operations. It is required for fueling the airplane if the fuel quantity indicators are to function.
2. Hydraulic-system test stand - This unit furnishes 25 gal/min, 3,000 psi hydraulic power for system preflight. Two units are required if the A and B hydraulic systems are to be preflighted simultaneously.
3. Ground air supply (start cart) - The aircraft requires 30 psi in the starter duct to start the first engine. Cross starting is possible after one engine is running, but this is not standard practice.
11.2 Fluid Requirements
Fuel MIL-T-5624, grade JP-5 or jet A per ASTM D1655
Oil Engine and auxiliary gearbox MIL-L-23699, Mobil Jet II
Hydraulic fluid MIL-H-5606 17 11.3 Airplane Grounding
The QSRA airplane must be grounded to earth during any ground service operation using the grounding jacks provided. For refueling, these jacks are located adjacent to both the single-point fueling station and the over-wing filler ports.
11.4 Servicing / Trained aircraft-service personnel will perform all maintenance, modifi- cations, and inspections required for the flight research program. Appropriate maintenance and inspection manuals will be available and used as reference material in performing this work. Complete up-to-date drawings (two sets) will be maintained in a location convenient to the airplane.
12. DOCUMENTATION LIST
Documents pertaining to the QSRA developed in support of contract NAS2-9081 are listed below. Distribution of these documents is controlled by the NASA Ames Research Center Quiet Short-Haul Aircraft Office.
Document Number
QSRA Configuration Definition Document NASA CR-152330 QSRA Maintenance Manual D340-14401 QSRA Operations Manual D340-13801 QSRA Predicted Performance D340-10100 Predicted Flight Characteristics D340-10500 QSRA Phase II Flight Simulation Mathematical Model NASA CR-152197
18 APPENDIX
LIST OF QSRA REPORTS
Boeing Commercial Airplane Company: Analysis of Contractor's Taxi and Flight Test of the QSRA. NASA CR-152322, 1980.
Boeing Commercial Airplane Company: Configuration and Definition. NASA CR-152330, 1981.
Boeing Commercial Airplane Company. The Development of a Quiet Short-Haul Research Aircraft - Final Report. NASA CR-152298, 1980.
Cochrane, John A.: Conceptual Studies of a Long-Range Transport with an Upper Surface Blowing Propulsive-Lift System. NASA TM-81196, May 1980.
Cochrane, J. A.; and Boissevain, A. G.: Quiet Short-Haul Research Aircraft — Current Status and Future Plans. AIAA Paper 78-1468, August 1978.
Cochrane, John A.; Riddle, Dennis W.; Stevens, Victor C.: QSRA — The First Three Years of Flight Research. AIAA Paper 81-2625, December 1981.
Eppel, Joseph C.: Mobile Automatic Test System (MATS) for Research Aircraft Instrumentation. Automatic Test Equipment Seminar Paper, June 1980.
Flora, Clarence C.; Middleton, Donald K.; and Schaer, Donald K.: Large-Scale Wind Tunnel Investigation of the Quiet Short-Haul Research Aircraft. Boeing Commercial Airplane Company. NASA CR-152203, 1979.
Flora, Clarence C.; Nicol, Laura E.; Marley, Arley C.; Middleton, Robin; Schaer, Donald K. ; Vincent, James H.: Quiet Short-Haul Research Aircraft Simulation Mathematical Model — Final Report. Boeing Commercial Aircraft Corporation. NASA CR-152197, 1979.
Hultman, Donald N.: Large-Scale Wind Tunnel Investigation for Future Modifi- cations to the Quiet Short-Haul Research Aircraft. Boeing Commercial Airplane Company. NASA CR-152349, 1980.
Kruppa, E. W.: Large-Scale Wind Tunnel Investigation of Quiet Short-Haul Research Aircraft (QSRA) Flightpath Control Concepts. Boeing Commercial Aircraft Corporation. NASA CR-152348, 1980.
McCracken, Robert C.: Quiet Short-Haul Research Aircraft Experimenters' Handbook. NASA TM-81162, January 1980.
McCracken, Robert C.: Quiet Short-Haul Research Aircraft Familiarization Document. NASA TM-81149, 1979.
19 Nickson, T. B.; Youth, S.; and Mlddleton, R.: Large-Scale Wind Tunnel Inves- tigation of the Quiet Short-Haul Research Aircraft. Boeing Commercial Airplane Company. NASA CR-152350, 1980.
Queen, S.; and Cochrane, J. A.: QSRA Joint Navy/NASA Sea Trails. AIAA Paper 81-0152, January 1981.
Riddle, Dennis W.: A Piloted Simulator Analysis of the Carrier Landing Capability of the Quiet Short-Haul Research Aircraft. NASA TM-78508, July 1980.
Riddle, D. W. ; Innis, R. C.; Martin, J. L.; and Cochrane, J. A.: Powered- Lift Takeoff Performance Characteristics Determined From Flight Test of the Quiet Short-Haul Research Aircraft (QSRA). AIAA/SETP/SFTE/SAE First Flight Testing Conference, November 1981.
Shovlin, Michael D.: Effect of Inlet-Airframe Integration on the Inlets of an Upper Surface Blowing Four Engine STOL Airplane. AIAA Paper 78-959, July 1978.
Shovlin, Michael D.: Interior and Exterior Fuselage Noise Measured on NASA's C-8A Augmentor Wing Jet STOL Research Aircraft. NASA TM X-73235, April 1977.
Shovlin, Michael D.: An Overview of the Quiet Short-Haul Research Aircraft Program. NASA TM-78545, 1978.
Shovlin, Michael D.: Upper Surface Blowing Noise of the NASA-Ames Quiet Short-Haul Research Aircraft. AIAA Paper 80-1064, June 1980.
Shovlin, Michael D.; Skavdahl, H.; and Harkonen, D. L.: Design and Perfor- mance of the Propulsion System for the Quiet Short-Haul Research Aircraft. AIAA Paper 79-1313, June 1979.
Vincent, James H.: Quiet Short-Haul Research Aircraft Phase 1 Flight Simu- lation Investigation, Vol. 1 Results: Flying Qualities and Flight Control System Design Requirements. Boeing Commercial Airplane Company. NASA CR-151964, 1977.
Vincent, James H.; and Marley, Arley C.: Quiet Short-Haul Research Aircraft Phase 1 Flight Simulation Investigation, Vol. 2 Mathematical Model. NASA CR-151965, 1977.
Wilcox, Darrell E.; and Cochrane, John A.: Quiet Propulsive Lift for Commuter Airlines. NASA TM-78596, June 1979.
Youth, Stan: USB Propulsion-Lift Aircraft Design and Performance Evaluation Study. Boeing Commercial Airplane Company. NASA CR-152387, 1980.
20 TABLE 1.- AIRCRAFT TOUCHDOWN LIMITS
The allowable sink speeds for landing are given as a function of landing weight in figure 4. Values for conditions specifically analyzed are given as follows:
Gross weight, Ib Allowable sink speed, fps
48,000 (or less) 12 55,730 11 GW >55,730 to 60,000 9 ZFW >42,000
Nose-wheel-first touchdown or pushing over abruptly into nose wheel is prohibited.
21 TABLE 2.- FLIGHT-TEST INSTRUMENTATION PARAMETERS - RMDU/A
Number of Parameter Comments measurements
Airplane environment Acceleration - C.G. 3 Lateral, longitudinal, normal Flightpath angles 3 Pitch, roll, heading Angle of attack 1 Nose boom Sideslip 1 Nose boom Altitude 3 Analog and digital transducer Airspeed 2 Nose boom Altitude rate 1 Radio altimeter Total temperature 1 Angular rates 3 Pitch, roll, yaw
Primary flight controls Control forces 3 Column, pedal, wheel Control positions 3 Column, pedal, wheel Surface positions 4 Elevator, aileron (2), rudder Spoiler position 4 Left and right inboard, outboard Flap positions 6 2 USB inboard, 2 USB center, 2 outboard Trim actuator position 2 Elevator, rudder
Automatic flight controls Servo command 3 Pitch, roll, yaw SAS actuator position 3 Pitch, roll, yaw
Fuel system Totalizer 4 Fuel used, engines 1, 2, 3, 4 Fuel temperature 4 Engines 1, 2, 3, 4 Fuel flow 4 Engines 1, 2, 3, 4
Propulsion Fan speed 4 Engines 1, 2, 3, 4 Core speed 4 Engines 1, 2, 3, 4 Total pressure (5.4), core exit 4 Engines 1, 2, 3, 4 Total pressure fan exit 4 Engines 1, 2, 3, 4 Total pressure fan exit 4 Engine 2, circumferential Fuel pressure, inlet 4 Engines 1, 2, 3, 4 Total temp. IGT (4.1), power turbine inlet temp. 4 Engines 1, 2, 3, 4 Power lever angle 4 Engines 1, 2, 3, 4
22 TABLE 2.- Concluded.
Number of Parameter Comments measurements
Boundary-layer control system Calibration duct dynamic pressure Engines 1, 4 Calibration duct static pressure 2 Engines 1, 4 Static pressure in duct. 2 2 aileron Refer to copilot's static pressure BLC duct total temperature Engines 1, 4
Miscellaneous Landing gear oleos 3 Right and left main, nose Event marker 1 Pilot's wheel Time - IRIG B 1 Slow code
23 TABLE 3.- FLIGHT-TEST INSTRUMENTATION PARAMETERS - RMDU/B
Number of Parameter Comments measurements
Airplane environment Acceleration, pilot's seat 2 Normal, lateral Wing vertical acceleration 2 Left wing, front and rear spar Accelerations, stabilizer 3 Left tip chordwise, left front spar and right front spar vertical Accelerations, vertical Lateral: Fin tip forward and stabilizer rear, midspan
Flight controls Flap position 6 All surfaces Spoiler position 4 All surfaces Flap actuator pressure feedback 6 Flap handle position 2 Outboard, USB Column force (CAS) 1 To SAS Speedbrake handle position 1 Wheel position (CAS) 1 Lateral flight-control actuator pressure
Hydraulic system Hydraulic pressure 2 Systems A and B Pump pressure 2 Engines 1 and 2 Suction pressure 1 System A
Propulsion Throttle positions (pilot) Engines 1, 2, 3, 4
Miscellaneous Fuel tank pressure Environment Control System (ECS) bleed air pressure 1 USB beep switch 2 Retract/extend - discrete USB flap handle 100° 1 USB flap handle not 0 1 Outboard flap handle not 0 1 DLC selected 1 Elevator out of trim 1 Fan exit temperature 4 Engines 1, 2, 3, 4 USB handle position 1 IRIG time 5 Seconds, minutes, hours
Engine overspeed trip Engine 1, 2, 3, 4
24 TABLE 3.- Concluded.
Number of Parameter Comments measurements
Electro-hydraulic actuator Position command Left and right inboard, out- board spoilers Inboard actuator current 1 Motor current 2 Left and right inboard spoilers Motor velocity 2 Left and right inboard spoilers Controller temperature 2 Left and right inboard spoilers Motor temperatures 4 Left and right inboard drives Inboard spoiler switch 1 Discrete 270 V d.c. 1 Discrete
25 4J
M O
cd a) CO 01
I 4-J M O JS DO
0) M 3
26 CM O u.
U * i*n N o CM " Ito ID u _l
0 3 CC rf X CC UJ U cc z CL < o Jr z a. z 1- >- Ul 0) u. CD
c o
UJ UJ UJ UJ UJ < 60 U m UJ •H O 1 0 O 0 0 0 CC 4-1 ia m D Z Z Z z Z < C Q O O
CM o CM CO to 00 O in ^ n 0 § 8 in 00 s cc III K C * _ bO Q. 4 •H ffl (> CO ^ Q c/> ac h- (U CO 00 O OJ Z z ICC cc rr cc C/) u. Q O UJ > UJ UJ Q X UJ 3 UJ -i _i POIL E c S CN CM o cc 0 o CM U) o o 0 o 0 0) 0 CO CO 0 o M 2 " 2 ? p o 1 1 3 2 i 00 N in c CO K o 0 o o o o 00 ^ o O M CM o ro N 1 1 in X D CO * s § s ? M S n n (M 0 1A O CO o f«k CM in in o 2 ^ | 1 O CO o in CO in o O o s S I1C * r- V) 3 •o C g u c* CC CO Q-" 0." z" UJ u P U) UJ cj o 2 ARE A (TRAP) , ft TAPE R RATI O T/ C BOD Y SIDE , % ASPEC T RATI O s CHOR D TIP , in . \- [ CHOR D ROOT , in' . | INCIDENCE , de g | DIHEDRAL , de g | VOLCOEFF V 27 200 r £ 192 £ 184 o o 176 168 (a) Roll inertia. 280 1:272 £264 o ro- £256 248 36 38 40 42 44 46 48 50 WEIGHT, 1000 Ib (b) Pitch inertia. Figure 3.- QSRA - Variation of inertias wi£h gross weight, 28 420 £ 412 U ~ 404 M 396 388 I (c) Yaw inertia. 26 r 25 o 24 o N -2 23 22 36 38 40 42 44 46 48 50 WEIGHT, 1000 Ib (d) Product of inertia. Figure 3.- Concluded. 29 OROSS 4000 Ib MAX. G.W. WING FUEL REMAINING 25 20 STRUCTURE + ENGINES DC « _ T> < LU ' ' 10 ^-^ ii i l l i i i i i i i i i III! 40 42 44 46 48 50 52 54 56 58 6 GROSS WEIGHT, 1000 Ib Figure 4.- QSRA weight and balance, main gear limit n. 30 INSTRUMENTATION ENGINE PRESSURE SENSORS M § 5 1 ' 33 in. r ' z AUX TANKS / 1 DO :n OC .00- / 1 INSTRU- O o AU>( TANKS ~—J MENTATION 0 L 1| '• - cc 1 1 1 1 o 1 T in i9 ^ O> Figure 5.- QSRA main cabin. 31 TAKEOFF GW = 50,000 Ib 2 U A = 50 O < u. -1 20 40 60 80 100 120 140 VELOCITY, knots CRUISE UH , ft/sec MAX ae MAX IDLE "/^^Txnt^ WINGi DESIGN LIMIT GW = 50,000 Ib 40 80 120 160 200 VELOCITY, knots STOL LANDING GW = 48,000 Ib 2 Ud =50 ft/sec 6 -MAX VDG cc USB > 30° O 1 U < u. § ° O -1 0 20 40 60 80 100 VELOCITY, knots Figure 6.- QSRA V-n diagrams. 32 ALL ENGINES OPERATING OUTBOARD ENGINE OUT 4000 WEIGHT, Ib 60,000 3000 45 39,000 LJJ 2000 1000 O 111 2400 WEIGHT, Ib 60,000 2000 £ uT O 1600 45 1200 O cc 800 Occ 400 66 70 74 78 82 86 90 94 UNCORRECTED FAN SPEED, % NL Figure 7.- Takeoff performance (outboard flaps at 59°, USB flaps 0°) 33 • OVER 35 ft OBSTACLE APPROACH • ACTUAL DISTANCE X 1.67 ANGLE, deg 2.5 6000 r n = SEA LEVEL n = 2500 ft n = 5000 ft 60 70 80 90 100 110 120 130 APPROACH VELOCITY, knots (a) Standard day. Figure 8.- Landing performance. 34 STANDARD DAY + 44°F OVER 35 ft OBSTACLE ACTUAL DISTANCE X 1.67 APPROACH ANGLE, deg 6000 2.6 5000 ( 4000 z 01 3000 u O Z 2000 n = SEA LEVEL n = 2500 ft n = 5000 ft 1000 50 60 70 80 90 100 110 120 130 APPROACH VELOCITY, knots (b) Hot day. Figure 8.- Continued. 35 / ' • OVER 35 ft OBSTACLE • STANDARD DAY 1000 7,deg UoJ . 2.5 <*" 500 oz 7.5 1o n = SEA LEVEL n = 5000 ft UJ ALL 7'S U 5 2 o *- o1 1 1 0 40 60 80 100 120 140 APPROACH VELOCITY, knots (c) Nonfactored distance. Figure 8.- Concluded. 36 CRUISE CONFIGURATION STANDARD DAY BLC SWITCHES OFF V = 130 KEAS, 88% NL .2 WEIGHT, Ib 50,000 , LU 5 45,000 iIj '' u '§ 20 u 10 00 0 1500 £ 1000 ui u. OQ u 500 10 15 20 ALTITUDE, 1000 ft (a) Time, distance, and fuel used. Figure 9.- Climb performance. 37 FLAPS RETRACTED - BLC SWITCHES OFF VCL = 130 knots ' SEA LEVEL STANDARD DAY SEA LEVEL HOT DAY (+44° F) WEIGHT, Ib 5000 40,000 4000 45,000 50,000 3000 2000 1000 c 1 0 «**--• CD S (b) Rate of climb - sea level. _l O LL O 111 5000 ft STANDARD DAY 5000 ft HOT DAY (+44° F) 2 5000 WEIGHT, Ib 4000 40,000 45,000 3000 50,000 2000 1000 64 68 72 76 80 84 88 92 96 FAN SPEED, % (c) Rate of climb - 5,000 ft. Figure 9.- Continued. 38 FLAPS RETRACTED - BLC SWITCHES OFF VCL= 130 knots 10,000 ft STANDARD DAY 5000 r 10,000 ft HOT DAY (+44° F) 4000 WEIGHT, Ib 40,000 3000 45,000 } 50,000 2000 1000 I£ 00 (d) Rate of climb - 10,000 ft. o Li. O 111 < 5000 15,000 ft STANDARD DAY cc 15,000 ft HOT DAY (+44° F) 4000 3000 WEIGHT, Ib 40,000 2000 45,000 50,000 1000 64 68 72 76 80 84 88 92 96 FAN SPEED, % (e) Rate of climb - 15,000 ft. Figure 9.- Concluded. 39 V = 130 knots NL = 50% STANDARD DAY W = 50,000 Ib 600 r- 10 16 ALTITUDE, 1000ft Figure 10.- QSRA descent capability. 40 30 sec, TURN = 28° n2 =1.13 1500 ft ALTITUDE V = 150 knots FLAPS UP DIMENSIONS IN n. mi. TAXIWAY T.O. wt = 45,000 Ib MISSION PROFILE THRUST TIME, min FUEL USED, Ib 1TO2 TAKEOFF, CLIMB TO 1500 150 ft ALTITUDE, ACCELERATE TO 150 knots, FLAPS UP 2 TO 3 FLY PATTERN TO FINAL 48% 3 2.7 248 APPROACH AT 1500 ft MAX ALTITUDE AND 150 knots 3 TO 4 DECELERATE TO 65.5 knots 30% 0.50 33 'MAX LOWER FLAPS TO 5USB = 60' 4 TO 5 APPROACH @ 7 =-7.5°, CL 43% F, 1.8 163 = 5.5 AND LAND APP 'MAX 5 TO 6 ROLLOUT AND TAXI BACK IDLE 2.2 93 7.95 687 Figure 11.- STOL mission cycle. 41 ACOUSTIC PANELS YF-102 ENGINE PRIMARY NOZZLE MIXED FLOW D-NOZZLE FAN BLEED / BLC BLC SYSTEM DUCT / EJECTOR AIR SUPPLY HIGH PRESSURE STARTER AIR SUPPLY HYD PUMP AND SERVO VALVE CONSTANT SPEED DRIVE GENERATOR CONSTANT SPEED DRIVE COOLER AND FAN Figure 12.- Engine installation. 42 1. FAN STAGE 6. CUSTOMER BLEED PORTS 2. FAN STATOR 7. COMBUSTOR 3. REDUCTION GEAR ASSEMBLY 8. GAS PRODUCER TURBINES 4. CORE AXIAL COMPRESSOR 9. POWER TURBINES 5. CORE CENTRIFUGAL COMPRESSOR 10. ACCESSORY GEARBOX 11. SUPERCHARGER (a) Cutaway. Figure 13.- QSRA engine. 43 1.442 (56.77) 0.667 * (26.25) n t I •- s ^^m J-h N ^ _ « . 1.0 T (40 t 0.471 (18.56) 5 Y f i r V I^M J Lp 1 NL 760C)(FAN) • 1"~~n (POVVEF*TUfRB IN E 17600) • Tr I1 IMH 196650 -- HJ.i SFC 0.41 MGT 169C)°F 921° C DIMENSIONS IN m(in.) MAX. THRUST 7500 Ib 33409.1 N WEIGHT DRY 1215 Ib 551.1 kg TOTAL AIRFLOW 267 Ib/sec 121.1 kg/sec CORE AIRFLOW 37 Ib/sec 16.8 kg/sec B.P.R. 6.2 (b) Geometry and uninstalled thrust. Figure 13.- Concluded. FIRE BOTTLES ENG ACCESSORIES DRAIN MAST FUEL PUMP STARTER FUEL CONTROL UN.T HYD PUMP CONSTANT SPEED DRIVE GENERATOR CONSTANT SPEED DRIVE COOLER AND FAN CONSTANT SPEED DRIVE ACCUMULATOR CONSTANT SPEED DRIVE FILTER (a) Layout. Figure 14.- QSRA nacelle. 45 NOSE COWL CORE COWL PRIMARY EXHAUST FAIRING PRIMARY NOZZLE ACCESSORY COOLING SYSTEM HP BLEED MANIFOLD ACCESSORY DRIVE SYSTEM (b) Engine buildup. Figure 14.- Concluded. C r 46 E S-100 120 35 90 100 x z cf 80 LU 80 LLI Q. CO 60 LLI 70 CC O 6o 40 2 50 20 CC cc 40 I 20 30 40 50 60 70 80 90 100 CORRECTED FAN SPEED, Np/^/6. % rpm (a) Fan speed, core speed, and thrust relationships. Figure 15.- YF-102 engine parameters. 47 SEA LEVEL STATIC 3% FAN BLEED 32 X HIGH SHAFT 7000 TORQUE LIMIT 6500 28 g DC X be 6000 x o HI I ^ 5500 24 5000 20 I- 4500 I I I I 20 40 60 80 100 120 AMBIENT TEMPERATURE, TAM/ °F -20 -10 10 20 30 40 50 °C (b) Installed thrust. Figure 15.- Concluded. 48 EXHAUST MIXING NOZZLE SHAPING PLUS 7 PNdB INLET NOISE 4 PNdB ATTENUATION WING SHIELDING DOUBLE LAYER LINING 12 PNdB AFT FAN ATTENUATION SINGLE LAYER LINING INNER AND OUTER FAN DUCT WALLS Figure 16.- Accoustic treatment in the QSRA nacelle. 49 r^ Ijj o co CO ^s O J-ol 4-> C Oo J-l QJ td 1 o M OJ Vi 00 •H 50 MIXING EJECTOR 2000 FAN z 400 AIR £ o 1600 AIR SUPPLY TO CD O BLC SYSTEM LL HIGH-PRESSURE ^300 ft AIR IN g 1200 CORE BLEED AIRFLOW cc CO UNREGULATED I D WITH 10% CORE FLOW 10% H- CC X 200 U H Z 800 o § oo o 1001 400 CO NORMAL REGULATOR VALVE OPERATION " °% LU MAX ENGINE THRUST I I I I I I J 30 40 50 60 70 80 90 100 ENGINE INSTALLED GROSS THRUST, FG~ percent Figure 18.- Net blowing momentum of a BLC nozzle. 51 S g IP s5||S*! 2 jj-t- ,J^ -3; Ol 1-m < §O1 Q FLIGHT-CONTROL t5* FUNCTION ^1- J e 01 C Ol < COLUMN] ]_ _ _ _1 PITCH CONTROL ^fc UJ>B 1 PITCH TRIM MODULATION! c " 1- 01 \- Ol ujO I >cc 1 10 PI LOT INPUTS | t |Sa3A31fr#'lff t SWITCH n I * 1 1 I- c x : _ . .. n A_ A -- A 1 14 ENGINE O | SH3indi/\ioosvs 3111OUHlllNOIllSOd 1 Cl b - llli i \jf«Jt^ zi n IB N "i \ 0 •i- < FLAP HANDLE H C 0 t * Q m c 35 01 Ol DISCRETES AILERON * SPOILER {OUTBD & USB) •• COMPUTERS SPOILERS ¥ 3 ft It- h SQUAT 0 g 52 m O L 0 SWITCH - —• 0 s 1 o b£ TO) OlT 1 OUTBD x5 55 ^ c ! — .tJ OUTBOARD S n " Ii 1 5 Z c oE L § J FLAP \ SO o|s L E tc_i5 * 2 SIS T COMPUTER 5 -> Q (AIR DATA INPUTS i g ---, 258 L-£2 .^ *— I nfe" 1— n ,.,_,. <2r^ 'F.il 1 _lZ L ^ATX a iz Slj S2£ 2§< 0 D 0' i- (< P t SvJ5 01 a. Q CO -1 O VS*^ '*!---* c O *^x> 1 * 0 c < co — 01 Z CO *t 01 LLbUI KIUAL . *_ 3 ROLL/YAW t- 2 X co — *• AIRPLANE MOTION j SAS COMPUTERS SENSORS L Ol o E S z Q 2. DIAGRAM SHOWS POWER-ON MODE ONLY H r-*1 \- C 3. CONTROL SURFACE SYMBOL INCLUDES POWER PEDAL|1 [ SWITCH A 1 DIRECTIONAL CONTROL AND SERIES ELECTRICAL FLIGHT- /\J? - 1 1 RUDDER CONTROL CONTROL AUGMENTATION ACTUATORS \f DIRECTIONAL TRIR -3 •H rj c O •J 2? s- • j •-1 —I S to u 1C 1) ELEVATOR AILERON FLAP (DOUBLE SLOTTED) SPOILERS RUDDER (DOUBLE HINGE) USB FLAPS FLAP (DOUBLE SLOTTED) AILERON SPOILERS (b) Flight-control surfaces. Figure 19.- Flight-control system. 53 '•O I 4-1 CO O 6 M CO 4-1 O O •H 13 CJ O H CO « h-5 ^ CO O tN (1) Vi 00 00 <£ 2 54 SPOILER 60 o> •8 •. 40 O i 20 -80 -60 -40 -20 0 20 40 60 80 WHEEL DEFLECTION, deg (b) Control gearing. Figure 20.- Continued. 55 1.2 1.0 .8 CM U •o .6 .4 MAX .2 10 20 30 40 50 60 70 80 WHEEL POSITION, deg (c) QSRA roll acceleration (CL = 4.6, weight = 48,000 Ib) Figure 20.- Concluded. 56 CO f>> CO O t-l *gJ O 00 I O •§ O •H PQ 0a0 01 57 GEARED BALANCE TAB A-A ELEVATOR AND GEARED TAB B -25 TRIM TAB +15 B-B ELEVATOR AND TRIM TAB (b) Elevator movement. Figure 21.- Continued. 58 TAKEOFF USB FLAPS = 0° 1.2 FC = 59° GW = 50,000 Ib 1.0 PM 1 .8 Z STOL LANDING O USBF = 66° FC = 59° D UJ .4 O CL. .2 60 70 80 90 100 110 120 V_, knots (c) Pitch acceleration. Figure 21.- Concluded. 59 CO >> 0) 4J c o M u 00 CO •H ctS o •H O 4-1 O a •aH IN (U >-l 00 •H 60 1.0 • .8 cc .6 in TAKEOFF _i UJ u g STOL LANDING >_j < h STOL APPROACH AT CL = 5.5 60 70 80 90 100 110 120 V, knots (b) Yaw acceleration. Figure 22.- Concluded. 61 111 o LLI cOc. u. o § 4-1 00 >> CO d o o p. cd i-0)l no •ri 62 LH POWER SYSTEM — H * L INBD I LEFT INBOARD SENSORS #2 - — I ACT J SURFACE • POSITION SPOILER A" HYDRAULIC SUPPLY 1/ CONTROL ^" • DISCRETE COMPUTER — »> * R INBD RIGHT INBOARD l APT SURFACE DIGITAL J ; COMPUTER SHUTDOWN 1 I DLC OUTPUTS ~1 If : . ' OUTBD LEFT OUTBOARD SENSORS #1 ACT SURFACE K SPOILER • POSITION > B" HYDRAULIC SUPPLY V CONTROL h" • DISCRETE COMPUTER R OUTBD RIGHT OUTBOARD APT "~ CIIDPAPP RH POWER SYSTEM * ACTUATORS TO BE DRIVEN BY SAMARIAN COBALT MOTORS Figure 24.- Electric spoiler command system. 63 ELECTRICAL POWER FROM LH SYSTEM LEFT i_cr i MMDUMnu i INBOARD ACTUATOR I FLAP SENSORS #2 M • POSITION USB FLAP CONTROL SOL 1 . „A" HYDRAULIC SUPPLY VALVE 1 ' • DISCRETE COMPUTER II RIGHT *-l RIGHT INBOARD INBOARD •« 1 ACTUATOR FLAP ELECTRICAL POWER FROM RH SYSTEM LEFT CENTER CENTER —I ACTUATOR FLAP SENSORS #1 II • POSITION USB FLAP CONTROL 1 SOL | , ,.B,, HYDRAULIC SUPPLY *l VALVE 1 ' B • DISCRETE COMPUTER M RIGHT CENTER 1 RIGHT CENTER ACTUATOR | ** =\ PI AP Figure 25.- Electric USB flap command system. 64 ELECTRICAL POWER FROM LH SYSTEM LEFT OUTBD LEFT FLAP -OUTBOARD ACTUATOR FLAP II SENSORS OUTBOARD "A" FLAP SOLENOID • POSITIONS CONTROL VALVE HYDRAULIC • DISCRETE COMPUTER SUPPLY M RIGHT OUTBD RIGHT FLAP OUTBOARD ACTUATOR FLAP Figure 26.- Electric outboard flap command system. 65 ANALOG PILOT'S AS NETWORK FTIS N u PANEL CIMTL RMDU 16 CHANNELS/BOX PNL AC A (8) BOXES 2G s" -^ ±5VDC \ V D X RC 1 D DIGITAL Tl ME °R NETWORK CC DE TAPE G U RCDR 1 C PANEL GENEf1ATOR C% T E 16 CHANNELS/BOX A R (1) BOX L S X ANTENNA \ r 1 i TM FILTER Y A* ANALOG L-BAND NETWORK RMDU XMITTER " PANEL B °G 50 CHANNELS/BOX l (4) BOXES Figure 27.- QSRA flight-test instrumentation system. 66 § 4J W >. W 0) 00 CN 60 •rl 67 a o •H 4-> 3 ,£> CO •H TJ I 3 00 •H 68 CM 4-1 CO co C8 U •H l-l 4J O 0) rH (1) CO (U VI 00 •H 69 LU a LU d 1- Iz- o >S9 < u cc ^o r<^~i^> T LU < O Q L >^o L-o£>—^ 1- ' , i Q ?» X * 1 UJQ r^i LU C^O a > LU -. Q > DC ^ M 1 DC ^ U LU LL 85 §3 <8 5 2 1 _l LU f" "" O n / cc i o< i 2 1 — .—J O r°~_"i 0 cc t_ CO I 1 0) ~~nrj] t t M i / 00 -iluJ ALU 1 •H LU 70 o to 0) CO C to CM 4J •H o (3 o o uO bO ^tt< m&i Oo-U w Q. C/J O 71 O O O O LIGHTING PANEL HYDRAULICS 1234 FIRE SWITCHES OUTBOARD FLAP USB FLAP LEVER LEVER\ ENVIRONMENTAL CONTROL USB PI TOT BEEP SWITCHES ENGINE SPEED BRAKE LEVER START THRUST LEVERS GUST LOCK LEVER (b) Overhead control panel. Figure 32.- Continued. 72 MARKER G/S f CONTROL PANEL IFF CONTROL DHC 60055-5 PANEL C6280A(P)/APX o o@o o ANTI-SKID 340-222028-1 340-222029-1 NO. 2 VHF NAV (VOR) CONT PANEL NO. 1 ADF REC C-3436A/ARN-30D CONTROL PANEL C-2275( )/ARN STATUS AND CONTROL 69 Y12847-1 (FLIGHT TEST) TACAN CONTROL PANEL C1763/ARN-21A O UHF CONTROL PANEL NO. 1 VHF NAV C-4677( )/ARC-51X (VOR) CONT PANEL C-3436A/ARN-30D VHF COMM CONTROL PANEL 614U-5 FLIGHT CONTROLS 340-2220181 340222019 1 ROLL SAS WHEEL STEERING INTERPHONE 999 INTERPHONE CONTROL PANEL CONTROL PANEL C-161K )/AIC( ) C-161K )/AIC( ) (PILOT) (COPILOT) J (c) Center pedestal. Figure 32.- Concluded. 73 LEFT AND RIGHT JETTISONABLE CABIN ESCAPE HATCHES PILOT'S HINGED OVERHEAD ESCAPE HATCH PRIMARY ENTRANCE BOARDING LADDER FLIGHT PI LOT AND COPILOT COMPARTMENT JETTISONABLE LOWER SIDE WINDOWS JETTISONABLE ESCAPE HATCH Figure 33.- Escape hatches. 74 1 Report No 2. Government Accession No. 3. Recipient's Catalog No NASA TM-81298 4 Title and Subtitle 5. Report Date QUIET SHORT-HAUL RESEARCH AIRCRAFT September 1981 FAMILIARIZATION DOCUMENT , REVISION 1 6. Performing Organization Code 7 Author(s) 8. Performing Organization Report No. Joseph C. Eppel A-8601 10. Work Unit No 9 Performing Organization Name and Address 532-02-11 Ames Research Center, NASA 11. Contract or Grant No. Moffett Field, Calif. 94035 13 Type of Report and Period Covered 12 Sponsoring Agency Name and Address Technical Memorandum National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D.C. 20546 15 Supplementary Notes This TM revises and updates TM-81149 by Robert C. McCracken published November 1979. 16 Abstract The design features and general characteristics of the NASA Quiet Short-Haul Research Aircraft are described. Aerodynamic characteristics and performance are discussed based on predictions and early flight-test data. Principle airplane systems, including the airborne data-acquisition system, are also described. The aircraft was designed and built to fulfill the need for a national research facility to explore the use of upper- surface-blowing propulsive-lift technology in providing short takeoff and landing capability, and perform advanced experiments in various technical disciplines such as aerodynamics, propulsion, stability and control, han- dling qualities, avionics and flight-control systems, trailing-vortex phenomena, acoustics, structure and loads, operating systems, human factors, and airworthiness/certification criteria. An unusually austere approach using experimental shop practices resulted in a low cost and high research capability. 17 Key Words (Suggested by Author (s)) 18. Distribution Statement Short-haul Aircraft design Unlimited Upper surface blowing Research aircraft STAR Category - 05 STOL 19. Security Classif (of this report) 20 Security Classif (of this page) 21 No of Pages 22. Price* Unclassified Unclassified 81 $9.50 'For sale by the National Technical Information Service, Springfield, Virginia 22161