An Overview of the Air Force`s Speed Agile Concept Demonstration

Transcript

51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition
07 - 10 January 2013, Grapevine (Dallas/Ft. Worth Region), Texas
AIAA 2013-1097
An Overview of the Air Force’s Speed Agile Concept
Demonstration Program
Downloaded by PENNSYLVANIA STATE UNIVERSITY on May 9, 2013 | http://arc.aiaa.org | DOI: 10.2514/6.2013-1097
Cale H. Zeune1
Air Force Research Laboratory, Wright Patterson AFB, OH, 45433
The United States Air Force faces numerous challenges in today’s changing world
environment, particularly in regards to its mobility fleet. To address these challenges the Air
Force Research Laboratory has been making focused investments in technologies for future
airlifters. A major emphasis by the Aerospace Systems Directorate has been on ‘speed
agility,’ a phrase coined for an aircraft that is capable of efficient flight at low speeds (~90
knots) and at transonic cruise speeds (Mach 0.8). A speed agile aircraft is able to takeoff and
land on short airfields as well as cruise faster than today’s military mobility fleet, giving it
unparalleled efficiency, access, and flexibility. In order to validate requisite technologies a
multi-year program was undertaken with government and industry partners that aimed to
refine host vehicle configurations, validate aerodynamic performance at low speed and at
transonic conditions, and develop and evaluate flight controls technologies to ensure safe,
robust aircraft operations. These goals were achieved for two separate hybrid powered lift
systems by way of testing at the NASA Langley 14’x22’, the AEDC National Full Scale
Aerodynamic Complex, the NASA Langley National Transonic Facility, the NASA Ames
Vertical Motion Simulator, and the AFRL LAMARS simulator.
Nomenclature
AEDC
AFC
AFRL
AJACS
BAA
BVWT
CBR
CCW
CFD
FOD
HPLS
HSPM
IBF
IML
LAMARS
LaRC
LM
ML/D
Arnold Engineering Development Center
Active Flow Control
Air Force Research Laboratory
Advanced Joint Air Combat System
Broad Agency Announcement
Boeing V/STOL Wind Tunnel
California Bearing Ratio
Circulation Control Wing
Computational Fluid Dynamics
Foreign Object and Debris
Hybrid Powered Lift System
High Speed Powered Model
Internally Blown Flap
Inner Mold Line
Large Ampl. Multi Mode Research Simul.
Langley Research Center
Lockheed Martin
Mach * Lift / Drag
NAART
NFAC
NTF
OML
PACAH
PSC
PSTF
PSP
QWSS
REN
SACD
STOL
TEMPO
TRL
USAF
USB
VG
VMS
North American Aviation Research Tunnel
National Full Scale Aero Complex
National Transonic Facility
Outer Mold Line
Pitch Attitude Command / Attitude Hold
Preferred System Concept
Propulsion Static Test Facility
Pressure Sensitive Paint
Quantitative Wake Survey System
Reversing Ejector Nozzle
Speed Agile Concept Demo
Short Takeoff and Landing
Transonic Effic. Mobility Planform Optimiz.
Technology Readiness Level
United States Air Force
Upper Surface Blowing
Vortex Generator
Vertical Motion Simulator
Introduction
The United States Air Force (USAF) is faced with a number of key challenges in today’s evolving world. An
uncertain future for oil production has resulted in volatile aviation fuel costs at a time when airlift requirements
continue to increase. A fragile global economy has led to a change in funding priorities by developed countries. The
1
Aerospace Engineer, Aerospace Systems Directorate; AIAA Membership pending
Distribution A: Cleared for Public Release; Case No. 88ABW-2012-6516; Ref. No. RQ-12-818
1
American Institute of Aeronautics and Astronautics
This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.
United States is easing out of the war on terror and confronting new global threats. Mobility resupply efforts being
conducted today typically use long, hard-surfaced runways to airlift in supplies, then load them onto trucks for final
delivery by a ground convoy; insurgent groups easily disrupt this predictable service by placing explosive devises
along roads and targeting aircraft with man portable air defense missiles. Basing and access are decreasing around
the world as sentiments toward the United States turn increasingly negative. The average age of the USAF’s
transport fleet is steadily increasing while an increased operational tempo is consuming the remaining useful
airframe life. Major acquisition efforts for recapitalizing the fleet are often behind schedule and over budget, while
others are cancelled or delayed due to issues with how the competitive award was conducted. The Air Force Special
Operations Command desperately needs to recapitalize and augment the capability of its aging C-130 gunship and
mobility fleets, but lacks the resources to do so. Still the special operations forces continue to play a preeminent role
in critical operations around the globe, often getting the job done unseen.
To address these challenges the Air Force Research Laboratory (AFRL) has been making focused investments in
technologies for future airlifters. A major emphasis by the Aerospace Systems Directorate has been on developing
‘speed agility,’ a phrase coined for an aircraft that is capable of efficient flight at low speeds (~90 knots) and at
transonic cruise speeds (Mach 0.8). The primary envisioned mission is to carry a 65,000 lb payload on a theater
resupply mission. A speed agile aircraft is able to takeoff and land on short, austere airfields as well as cruise faster
than today’s military mobility fleet, giving it unparalleled efficiency, access, and flexibility.
Background
In August of 2007 AFRL issued a request for proposals to validate the development of key speed agile
technologies. A number of lower technology readiness level (TRL) development efforts had been going on for some
time, mainly consisting of smaller scale proof of concept wind tunnel tests or component technologies and host
aircraft configurations [1]. This subsequent validation program was called the Speed Agile Concept Demonstration
(SACD) and had the goal of achieving a TRL of at least five (5) by 2010 on an integrated mobility configuration in
the areas of high lift, efficient transonic cruise, and flight controls, in order to support future technology
development and acquisition activities. Specific objectives were configuration refinement, low speed performance
validation, transonic cruise performance validation, and flight controls development and handling qualities
evaluations. An increase in planned payload weight for the host vehicle (to 65,000 pounds) and cross sectional size
over previous concept vehicles had led to the need for a vehicle refinement. The SACD objective vehicle is to be
capable of carrying the 65,000 lb payload a radius of 500 nm, landing on a 2,000 ft, California Bearing Ratio (CBR)
5 site, takeoff with the same payload, and return 500 nm; it should also be able to self-deploy with some small
payload a range of 3,300 nm at cruise speeds greater than Mach 0.8.
These mission capabilities were specified to ensure that the
lift system and underlying technologies were being developed
in the proper capability context and are not meant to be official
aircraft requirements or imply that the SACD program is an
aircraft development program. Indeed the intent of this program
has been to mature the underlying components and technologies
to a point where they are proven feasible and effective for use in
a variety of aircraft configurations and can be utilized to satisfy
a wide range of future Air Force requirements. The integrations
and vehicle configurations currently in development are one
particular instantiation devised to be particularly challenging in
order to push technology toward enabling greater capabilities in
theater airlift. As additional capabilities or constraints are
imposed on the aircraft or removed from it the objective aircraft
design will change to accommodate accordingly, but it is
envisioned that the core technology components will be
adaptable to these configuration changes and still very relevant
and applicable in a different form of the integration.
To meet these objectives, AFRL awarded two efforts to
separate industry teams to mature two different implementations
of a hybrid powered lift system (HPLS). One approach pursued
by The Boeing Company was a hybrid powered lift system Figure 1 – Host vehicle configurations:
comprised of an upper surface blowing (USB) system inboard Lockheed Martin (top) and Boeing (bottom)
2
and circulation control / internally blown flaps scheme on the outboard wing sections. A second approach,
developed by Lockheed Martin, uses circulation control / internally blown flaps on the outboard wing sections and a
novel ejecting, reversing nozzle inboard. These vehicle concepts are shown in Figure 1. In the course of maturing
both of these lift system concepts and their integration into representative host vehicles, a series of design and
integration activities, low speed wind tunnel tests, transonic wind tunnel tests, and flight control simulation activities
were conducted. A brief overview of these technical activities is presented in subsequent sections.
Boeing Developmental Program
AFRL entered into a cooperative agreement with Boeing to mature Speed Agile technologies in early 2008.
Under this agreement, both parties allocated funds and other in-kind resources to the effort. Boeing had been
working in this technology area preceding contract award and brought those technical efforts into the program.
AFRL, over the course of the contracted effort, provided several government test facilities to the Boeing team as
their share of the partnership. By early 2008 Boeing had already developed a baseline vehicle concept and an
integrated lift system package, and built and tested a small scale low speed wind tunnel model at their short takeoff
and landing (STOL) tunnel in Philadelphia. This vehicle configuration featured a cargo box and fuselage width
somewhat smaller than the baseline characteristics requested by AFRL in their Broad Agency Announcement
(BAA) for Speed Agile.
A. Configuration Development and Lift System Integration
The purpose of this first task was to refine an integrated mobility configuration (preferred system concept – PSC)
with engine company support, incorporating advancements in the high lift for STOL performance, and aerodynamic
design for efficient transonic cruise. The first subtask was to update the configuration using flight and ground
operational mission and utility guidelines derived from the AFRL Speed Agile Broad Agency Announcement. The
second subtask was to refine the concept using the latest aerodynamic and performance data available. The concept
refinement included a preliminary structural definition and blowing subsystem design. The third subtask was the
engine company design studies, which included engine bleed air design and integrated thrust reverser and nozzle
design.
Immediately after the program start, the Boeing proposal concept was updated to the larger payload bay in the
BAA. This concept also included the wing planform from a previous study (Transonic Efficient Mobility Planform
Optimization – TEMPO) which optimized the aerodynamic shape to reduce transonic drag. Through several design
iterations the concept was improved. As the concept matured, following versions had a revised airfoil and planform
to match the lines of the final low speed wind tunnel model, a revised inlet, as well as a new nozzle and reduced
chord Upper Surface Blowing (USB) flap. This configuration also had improved landing gear definition and
integration. The final iteration on the vehicle matched the lines of the low speed and transonic wind tunnel test
models, and had a modified tail and inboard wing leading edge.
Through cooperation with an engine company partner, a number of propulsion related studies were conducted.
An assessment of the baseline nozzle and an alternate nozzle were accomplished, and resulted in recommendations
to improve nozzle performance. In order to mature a thrust reverser concept a series of iterations involving design
and computational fluid dynamics (CFD) analyses were performed. Lastly, the development of a bleed air extraction
scheme to supply air to the outboard blown flaps was undertaken. Subsequent work planned for 2009 included
improved foreign objects and debris (FOD) tolerance.
Some portions of this task were terminated in December 2008 in order to preserve resources for low speed
performance validation and flight control system development and assessment tasks. The PSC aero, weights, and
performance analyses were finalized. Analyses included a weight breakdown, high speed lift and drag, low speed lift
and drag (including powered lift), installed engine thrust and fuel flow, mission, takeoff, and landing performance.
The aero performance analysis done on the resulting concept showed it was capable of efficient Mach 0.8 cruise
using the refined geometry and aerodynamic results from TEMPO. When sized to the missions/payloads specified in
the BAA, this concept was capable of assault takeoff and landings from a 2,000 ft field at mid-mission weight.
B. Wind Tunnel Testing
The wind tunnel test effort in support of the Speed Agile program consisted of a series of low speed, high speed,
and static nozzle rig test entries that spanned from July 2007 to October 2009 and included a total of 2,278 test
facility occupancy hours. The test program included two 1.27% scale low speed unpowered models, two 5% scale
low speed powered models, a USB static development test rig, and a 3% scale transonic unpowered model. The low
speed tests included both the initial narrow body configuration as well as the revised wide-body configuration; the
3
high speed model represented only the revised wide-body configuration. A timeline of the wind tunnel test program
is shown below in Figure 2.
1. Low Speed, Full Configuration Test Entries (LB-638A & LB-638B)
The first two low speed test entries consisted of the initial narrow-body 5% scale full-span powered model
designated as LB-638. The model was first tested in the Boeing BVWT 20’x20’ tunnel (the “A” entry) which was
completed in August of 2007. This test entry was used for the initial development of the active flow control (AFC)
system in the wingbox and trailing edge flaps, assessment of the USB system turning performance, and the
acquisition of the initial low speed performance database. The test results indicated that the AFC system worked
well and provided large improvements to the high lift system performance, but the USB flaps only generated 30° of
flow turning which was well short of the turning requirement determined from the preliminary performance
estimates.
Figure 2 – Boeing wind tunnel test program timeline
The second low speed test entry (designated LB-638B) of the same 5% scale model was done in the National
Aeronautics and Space Administration Langley Research Center (NASA LaRC) 14’x22’ wind tunnel. The same
powered model was again tested over a range of thrust settings and the test objectives included refinements to the
AFC system and the USB/nozzle configuration. Further objectives included demonstration of the powered full-span
model test capability in the 14’x22’ facility, acquisition of on-body and off-body flow visualization via fluorescent
mini-tufts and the Boeing Quantitative Wake Survey System (QWSS), and acquisition of initial control surface
effectiveness and lateral-directional characteristics. The refinements to the AFC system included narrower AFC
slots on the flaps, alternate chordwise flap slot locations, and the evaluation of pulsed AFC air supply. All test
objectives were met, however the simple refinements to the USB system did not perform as well as hoped. It became
clear at this point that a dedicated nozzle / USB development effort would be necessary to meet the overall Speed
Agile program requirements of demonstrating a TRL of 5.
2. 1.27% Scale Unpowered Low Speed Tests
Two small scale full-span unpowered wind tunnel models were built and tested in the Boeing North American
Aviation Research Tunnel (NAART) test facility during the second quarter of 2008. The first test entry (NAART
147) represented the narrow-body configuration and the second entry (NAART 149) represented the refined widebody configuration. Both models included slats and trailing edge flaps as well as a simplified representation of the
USB flaps since the model was unpowered. Both test entries were used for the parametric development of the wing
lateral-directional control surfaces as well as parametric evaluation of the tail and ruddervator control effectiveness.
The results from these tests, combined with the results from the larger scale 5% LB-638A and B test entries, were
4
used to build the initial vehicle math model for the basic longitudinal and lateral-directional stability characteristics.
Furthermore the control surface database for the most promising control effectors was used for down selecting to the
control surfaces defined for the final 5% scale wide-body wind tunnel model which was scheduled for testing after
the USB / nozzle static rig development was completed.
3. USB Static Development Tests (LB-639A, B, C)
After completion of the LB-638B powered
test entry in the LaRC 14x22 facility it was clear
that a dedicated nozzle / USB development test
program would be necessary to achieve the flow
turning goal. A static rig development model
was designed and built which represented only
the inboard trailing edge portion of the vehicle
configuration consisting of the nozzle and USB
flap (depicted in Figure 3). The model was
designed to support a parametric study of the
nozzle configuration as well as different USB
flap configurations. The model was tested in the
Boeing Propulsion Static Test Facility (PSTF) at
St. Louis in three separate entries (LB-639A, B,
and C) during 2008. Various nozzle
Figure 3 – USB Nozzle and Flap Static Rig Model
configurations including thrust deflecting flaps,
nozzle kick-down flaps, and nozzle kick-out flaps were evaluated as well as a range of USB constant radius and
spiral radius flap shapes. The final USB configuration had a considerably more aggressive geometry than heritage
USB designs, yet was able to produce a net turning angle change from cruise to landing of 51 degrees. Incorporating
a high-performance spiral ‘conformal’ flap coupled with nozzle changes, and vortex generators enabled the aircraft
configuration to meet its landing field requirement. This goal was achieved using an identical aircraft configuration
to that which meets efficient transonic cruise requirements. The final optimized nozzle / USB flap configuration
developed during these tests were incorporated into the final wide-body LB-640 low speed model. For more
technical information on this activity, please see Reference [2].
4. Final Low Speed Powered Test (LB-640A)
All of the knowledge gained from the prior LB-638, LB-639, and NAART wind tunnel tests were incorporated
into the design of the updated wide-body LB-640 model. This full-span powered model included the refined AFCenabled high lift system, the revised USB nozzle and flap, and a full set of control surfaces. The model is shown in
Figure 4, with the hybrid powered lift system components called out. The outer mold lines of this model were
consistent with those of the following transonic test model. The LB-640 model was tested in the NASA LaRC
Figure 4 – LB-640 Model, with upper surface blowing and internally blown flaps
5
14’x22’ facility in 2009 which was by far the largest test entry of the Boeing Speed Agile test program consisting of
1500 data runs acquired during 1100 hours of testing. The test objectives were split into two phases: The first phase
consisted of the validation of the re-designed USB / nozzle configuration, further optimization of the AFC high lift
system, and the acquisition of the final wide-body aero data for performance. The second phase, which was the bulk
of the testing, consisted of the acquisition of the aero data required for the flight simulator database including all
cruise, takeoff, and landing configurations as well as the evaluation of control surface effectiveness, failure modes,
longitudinal trim requirements, and lateral-directional characteristics.
The test results showed that the re-designed USB / nozzle configuration met the flow turning goal, the AFCenabled high lift system was validated, and the final corrected trimmed lift and drag characteristics were adequate to
meet the takeoff and landing field length requirements specified for the Speed Agile program. Furthermore the
database acquired during the test supported the high fidelity, low speed simulation effort and that the control
surfaces were adequate to meet one-engine inoperative trim, roll requirements, and trimmed sideslip condition.
5. High Speed NTF Test Entry (LB-641A)
The final wind tunnel test consisted of a 3%
scale full-span unpowered model built under a
cooperative agreement with NASA for testing in
the LaRC National Transonic Facility (NTF).
The outer mold lines of this model were
consistent with the LB-640 final low speed
model. A picture of the wind tunnel model is
shown in Figure 5 as installed in the tunnel’s
test section. The entry in the NTF was
completed in October of 2009 and the results
indicated TRL validation of the high speed
design including the cruise Mach number
requirement of M = 0.80. The data from this test
entry was used to support the database required
Figure 5 – The Boeing configuration installed in the National
for the high speed portion of the simulation
Transonic Facility
effort.
C. Flight Controls Development and Handling Qualities Simulation
A six degree-of-freedom simulation and closed-loop flight control system including representative control laws
was developed and a formal piloted evaluation of this simulation took place from 31 March to 15 April 2011 at the
Vertical Motion Simulator Facility (VMS), NASA Ames Research Center, Moffett Field, CA. A total of 8 pilots
participated in the experiment: 5 test pilots, 3 fleet-qualified operational medium mobility pilots, and one with
combined operational and flight test experience. The overall handling qualities of the aircraft were evaluated during
a wide range of maneuvers, including terminal and non-terminal maneuvering. The primary objective was to
demonstrate Level 1 flying qualities during nominal terminal and non-terminal area operations, and Level 2/3 flying
qualities for degraded modes of operation in terminal area operations. In general, airplane characteristics were
described as Level I during nominal operations, particularly after the introduction of Pitch Attitude Command
/Attitude Hold (PACAH) mode for approaches and landings. Off-nominal operations were frequently described as
Level II; however, the off-nominal conditions (e.g. crosswind, lateral offset) were extremely demanding, often
approaching and occasionally exceeding the flight envelope. In addition, takeoffs and landings were intentionally
conducted assuming an operational environment that allowed no margins for engine or subsystem failures during
flight (i.e. “Assault Rules”).
Initial time-domain modeling of the airplane model highlighted several design shortcomings, including
inadequate directional control power in the approach configuration, and inadequate excess power during go-arounds.
These deficiencies were accommodated through conceptual design changes that must be addressed in a detailed
design.
Several flight controls deficiencies were also identified, in particular inadequate speed control in Angle of Attack
(AOA) command mode, pitch control unpredictability and sluggishness on takeoff and approach, and high workload
in the directional axis during offset approaches and engine failures. The AOA command deficiency was addressed
with the incorporation of PACAH mode, which highlights the benefits of a rapidly-reconfigurable simulation, and
the remaining deficiencies have been noted and flight control changes for improvement have been identified and
implemented in a post-test simulation update that has not been fully evaluated.
6
Within the scope of these tests the handling qualities of the MMT2-100 configuration are acceptable to
satisfactory under all conditions evaluated when operational necessity, pilot training, and operational procedures
refinement are considered. The control architecture and allocation scheme exhibit excellent potential for use in a
Speed Agile platform or highly unstable large aircraft regardless of external configuration. Demonstration of
acceptable handling qualities up to the limits of the flight envelope in a high-fidelity motion-based simulation
representative of the flight environment with operators in the loop is indicative of a TRL of 5 (Component and/or
breadboard validation in relevant environment). The reader is encouraged to consult Reference [3] for a more
detailed presentation of the aerodynamic model and flight control system developed and evaluated on this program.
Lockheed Martin Development Program
AFRL awarded a contract to Lockheed Martin (LM) in early 2008 to validate the technologies they had been
developing in the recent past in the Speed Agile realm. Lockheed Martin’s vehicle concept was built around a hybrid
powered lift system composed of two pieces: inboard, using the main exhaust stream, is a novel mechanism called
the reversing ejector nozzle (REN) that operates in several modes and provides thrust vectoring, thrust
augmentation, and reversing functions; outboard, on the main portion of the trapezoidal wing, is an internally blown
flap. Over the course of five years, a number of tasks were completed to integrate and validate the lift system, the
transonic aerodynamic design, and the flight controls to a TRL of 5. These tasks are described below.
A. Lift System Concept
The Lockheed Martin configured hybrid powered lift system is built around a reversing ejector nozzle (REN);
this innovative nozzle mechanism is depicted in Figure 6. The REN operates in three different modes and has
several features that are of particular interest. In high lift mode at low speeds an upper door is opened, allowing air
from the upper surface to be entrained with the main exhaust flow, and the thrust is augmented as the system
functions like an ejector. Lift is also augmented as the entrained air over the upper surface creates additional lift on
that surface. The pitching moment change is minimized since the thrust post is acting very near the nozzle exhaust
plane and not appreciably aft, as in some systems – this allows the aircraft to have a smaller tail than would
otherwise be necessary to trim the pitching moments in powered lift mode. For deceleration on the ground a
“bucket” device is engaged, directing the main exhaust flow up and forward over the upper surface of the wing,
functioning as a thrust reverser. For up and away cruise flight the entire system stows for efficient cruise thrust
production.
For lift augmentation on the outboard, traditional portion of the wing, LM has developed a unique approach to an
internally blown flap architecture that uses bypass stream air extracted from the engine, routed out to a plenum
behind the aft spar that exhausts through a slot near the knee of the flap. At low speeds this mechanism provides
aerodynamic lift augmentation by controlling separation off of a deflected aft flap at low blowing rates, and can
provide super-circulation lift at higher blowing rates. Another unconventional use of the blown flaps is at transonic
cruise conditions to reduce compressibility drag; though the blowing rates (based on momentum) are very small, this
method has shown potential (through computational studies) to alter the shock location and beneficially affect drag.
B. Configuration Refinement
This task involved the development of the
integrating concept configuration that meets
all mission, payload, and military utility
requirements. As mentioned previously, the
Speed Agile mission ‘desirements’ were
updated to include a heavier payload carriage
and a larger cargo box size; these changes
necessitated a configuration change to a four
engine airplane, which in turn required reintegration of the integrated high lift system
to meet the high lift generation and cruise
drag characteristics. Major activities that
were performed under this task included
configuration sizing and loft, detailed
aerodynamic analyses using Navier Stokes
CFD to meet cruise ML/D goals, propulsion Figure 6 – Lockheed Martin reversing ejector nozzle mechanism
7
integration, tail arrangement and sizing, internal systems layout, configuration documentation, a preliminary
assessment of lateral directional control and handling qualities, aft end ramp integration, and conceptual layout of
major structural members. A more thorough discussion of this process is described in Reference [4].
One task that proved particularly challenging was the integration of the hybrid powered lift system, due to the
competing requirements to minimize cross sectional area and reduce transonic drag, accommodations for structural
and air duct pass-through, and propulsion integration concerns that all happen near the wing root. The use of
advanced CFD automation tools and shape design utilities were critical in finding an acceptable solution to the
integration problems encountered. For example, the design concept for supplying air to the outboard blown flaps
was based on using cooler fan bypass air. This flow stream is at relatively low pressure, and to transport the required
amount of mass with low losses (at relatively low internal flow speeds), larger area ducts were needed. These ducts
take air out of the fan stream, diffuse it, turn the flow outboard, and feed into a larger plenum aft of the rear spar. All
of this routing must stay in a compact space, clear of main structure, and make benign turns to minimize duct losses.
This is also a key structural area due to the fact that wing bending is large, high lift and thrust loads are concentrated
here, and space must be left for actuation and mechanization of the reversing ejector nozzle. CFD tools were used to
analyze and optimize the OML and IML with the goals of maximizing lift and minimizing drag and internal losses,
both at low and high speeds.
After many months of design, refinement, and analysis a configuration was developed that met mission
performance goals, including cruise speed and efficiency, short field performance, stability and control and handling
qualities characteristics, and other military utility factors. Final transonic aerodynamic performance came in at the
goal levels, namely an ML/D max of over 13 occurring at Mach numbers greater than 0.8. The goals for short field
performance were also achieved, yielding a configuration capable of taking off and landing in less than 2,000 ft
(over a 50 foot obstacle) while carrying 65,000 lb.
C. Wind Tunnel Testing
1. Low Speed Performance Validation
A 23% scale wind tunnel model was designed based on the full scale preferred system concept for testing at the
National Full Scale Aerodynamics Complex (NFAC) at Moffett Field, CA. This tunnel is operated by the Air
Force’s Arnold Engineering Development Center and has two test sections; the larger section is 80’x120’ and has a
maximum flow speed of 100 kts, while the smaller section is 40’x80’ and has a maximum flow speed of 300 kts.
Testing at this facility allowed the technology to be validated in a relevant environment at large scale using turbofan
engines to test the lift system’s architecture using hot flows.
The wind tunnel model was congruently scaled geometry of
the full scale objective vehicle, except for the inlets and
number of engines; Williams International supplied FJ44
engines for testing and to reduce risk to the engines a set of
conventional bellmouth inlets were used. Two FJ44 engines
were selected to power the model because it provided the
appropriate thrust levels for thrust scaling and had the right fan
pressure ratio to supply fan bypass air to the blown flaps. The
model featured many actuated and remotely controlled control
surfaces to aid in test productivity. The hybrid powered lift
system was able to be tested in all three of its modes: clean
cruise, high lift, and reversing. Testing was conducted from
March through October of 2011. The initial entry into the
80’x120’ tunnel ran from March through May, and the later
entry from August through October. The main objectives of
the test were to establish a TRL of 5 for the lift system concept
by validating the architecture with hot flows, comparing
predicted levels of aerodynamic performance at low speed to
those measured in the wind tunnel, and to gather an
aerodynamic database for follow-on flight controls and
handling quality simulations.
The 80’x120’ entry was intended to provide the static
thrust calibrations in cruise, reverse, and high lift modes, as Figure 7 – Lockheed Martin Low Speed
well as wind on data for each at wind speeds below 100 kts. Powered Lift Model installed in the NFAC
The reverser mode wind on data was not obtained due to 80'x120'
8
engine surging that was observed during the static testing at intermediate engine power. The wind on data in this
tunnel also provided a basis for assessing the wall correction blockage affects in identical runs in the 40’x80’ tunnel
later. The model in shown installed in the 80’x120’ in Figure 7.
The 40’x80’ entry was where the principal high lift configuration, wind on data was obtained. The 40’x80’
testing acquired basic aerodynamic performance as well as control surface perturbation data to assess incremental
effects. Additionally control surface effectiveness and stability data was acquired to develop a database for the flight
control and simulation tasks. Finally, similar cruise configuration control surface effectiveness and stability data was
planned to be acquired but due to a model failure only an unpowered series was executed with the control surfaces
in neutral positions.
The test finished in October, successfully fulfilling the primary objectives. High lift performance was excellent,
achieving goal maximum lift coefficient levels and comparing favorably with pre test CFD predictions. Valuable
control surface increments and failure mode data were gathered, and the lift system performed very well under
representative thermal conditions. Many lessons on testing large scale powered lift models were gleaned, including
wall correction methodologies, instrumentation best practices, ventilation requirements, fuel handling processes,
model construction, and operation of live engines inside the wind tunnel. More work will need to be done to
understand the short comings of the reversing mode as tested. The model is now in storage and can be used for
future testing, pending some relatively minor repairs.
Further details regarding the test and associated CFD activities can be found in References [5] and [6].
2. Transonic Cruise Performance Validation
In order to validate the aerodynamic performance of the vehicle’s shape design and to test the effectiveness of
the blown flaps at reducing drag during cruise conditions, a semispan jet effects model was built and tested at NASA
Langley in the National Transonic Facility. The Speed Agile High Speed Powered Model (HSPM) wind tunnel test
was conducted at the NASA Langley Research Center National Transonic Facility (NTF) wind tunnel from 31 May
2011 through 24 August 2011; an image of the model is shown in Figure 8. The HSPM was a 5.25% scale semi-span
model and was delivered to NASA/Langley NTF on 31 May 2011 by the model vendor Advanced Technologies
Incorporated. During the month of June the model was installed in the NTF test section, and several weeks of
instrumentation hookup and checkout followed. Preliminary balance-only data were obtained in early July. The
majority of air-on testing occurred during July and August. During this period, there were approximately 482 windon runs acquired, with 290 runs in air (120 deg-F) and 192 runs in cryogenic mode (-50 deg-F). The 192 runs in
cryogenic conditions include 31 runs of cryogenic Pressure Sensitive Paint (PSP) data. On 18-19 August, a series of
20 runs were accomplished to measure the wake behind the main engine nozzle and the blown flaps slot. Model
removal was scheduled for late August. After the model removal, NASA conducted additional balance calibration
work to help resolve balance issues discovered during testing.
The majority of the planned test matrix was completed during the 380 hour test. Model deformation
measurements were made throughout the test using the NTF Visual Model Deformation (VMD) camera system.
Pressure Sensitive Paint (PSP) testing was accomplished in cryogenic mode. The wake survey done at the end of the
test was not part of the original test schedule.
These data have provided useful information
concerning the main nozzle and blown flap
slot exit conditions.
Several significant problems were
encountered during the test that ultimately
affected the team’s ability to accomplish
several of the test objectives with suitable
confidence in the resulting data. The
sidewall balance used for this test had an
unanticipated yet significant sensitivity to
both bellows pressure and bellows
temperature. Changes made to the sidewall
mounting system to accommodate the air
station resulted in an inability to adequately
control balance temperatures within the
necessary
margin,
especially during
cryogenic runs. Corrections for these effects Figure 8 – Lockheed Martin transonic wind tunnel test model
were attempted following the test, but they installed in the NTF
9
failed to suitably reduce scatter in the data to the point where it was useful for validation. Other minor problems
were encountered with the pressurized air supply valve (which controls the mass flow through the model main
nozzle), and with the Electronic Scanning Pressure (ESP) modules, though these were eventually overcome.
Throughout the test there were problems with losing filler material used to fill the model’s external fastener holes,
though this effect on the resulting drag data is believed to be negligible. Early in the test, one of the four total
pressure rakes located at the model main nozzle exit was lost, as well as the thermocouple used to measure
temperature in the main nozzle duct; additionally, one pressure tube from one of the blown flap slot total pressure
rakes was blown out of the model. Due to redundant instrumentation schemes, these issues were mere annoyances
and the integrity of the internal pressure diagnostic measurements were adequately maintained.
Using suitably corrected data taken for the unpowered air cases, the basic shape design of the aircraft was
validated, showing a peak ML/D value at Mach 0.81. The associated ML/D values were questionable and a peak
value cannot be conclusively stated at this point. Similarly, the aero performance of the powered cases in air,
powered and unpowered runs during cryogenic conditions, and the effect of IBF slot blowing at transonic conditions
cannot be considered conclusively validated at this point. It should be noted that this was the first production wind
tunnel test at the NTF utilizing the newly developed capability to test semispan models transonically at high
Reynolds numbers and the ability to supply pressurized air through the balance center to simulate propulsion effects
and flow control. NASA is working to resolve issues encountered and has conducted a test of a research model to
demonstrate suitability for future testing. NASA and AFRL are exploring options for re-testing this Speed Agile
model at a future date. The Speed Agile program is indebted to the team at NASA who invested in, developed,
tested, and operated this new tunnel capability.
A discussion of the transonic design of the vehicle and this transonic testing is expounded on in Reference [4].
D. Flight Controls Development and Handling Qualities Simulation
In early November 2012 a team of simulation and flight control engineers from the Lockheed Martin Speed
Agile program, along with 9 pilots, and the staff of AFRL/RQCD successfully completed a handling qualities
simulation in the Large Amplitude Multi-Mode Aerospace Research Simulator (LAMARS) facility at Wright
Patterson Air Force Base in Ohio. This was the last major experiment in the LM Speed Agile program. Its objective
was to evaluate (using the Cooper Harper rating system) the flying qualities of a STOL aircraft equipped with a
novel hybrid powered lift system. This particular simulation was populated with aerodynamic data gathered in the
preceding wind tunnel test tasks (largely the AEDC NFAC). Nine pilots, including C-130 and C-17 pilots
participated from Air National Guard, Lockheed Martin, 445th Airlift Wing, and other reserve units. Evaluations
focused largely on approach to landing, takeoff, and attitude capture tasks. Overall, preliminary results were
favorable.
The LAMARS is a five degree of freedom simulator capable of several g’s of horizontal and vertical
acceleration, as well as 25 degrees of rotation in roll, pitch, and yaw. A single pilot simulator, it is made up of a
large, hemispherical dome mounted on the end of a 20 foot cantilevered arm and actuated by a hydraulic system.
During this set of simulations, a center stick was used for primary control; generic heads up and heads down
displays were configured to inform the pilot during the evaluations. Takeoff, approach, and landing tasks were
conducted at two sites (Los Angeles International airport on a standard day and Kern Valley/Isabella Lake airport on
a hot day) to offer a variety of visual and density altitude conditions.
A variety of tasks were accomplished by the pilots, in increasing levels of complexity/workload. The first major
block of tasks were powered approach mode maneuvers, including trim, bank capture, glideslope capture, angle of
attack capture, and velocity and heading changes. Next came straight in approaches with calm winds, gusts, and
crosswind conditions; takeoff tasks were also conducted in a variety of wind conditions at both sites. Following the
straight in approaches, glideslope and localizer offset approaches were conducted, occasionally with a failed engine.
Finally, a variety of more operationally relevant approaches were evaluated at Kern Valley under a sampling of
wind conditions, including dog leg approaches, full pattern approaches, high speed and low speed approaches, and
steep approaches. Hot day takeoff tasks were performed with wind varied and engines inoperative.
Pilots evaluated the aircraft’s handling qualities using the Cooper Harper rating scale, and also evaluated the
tendency of the airplane to enter into pilot induced oscillations. Final detailed conclusions relating to handling
qualities are still being processed, though initial impressions were positive in general. Pilots offered excellent
feedback on the aircraft’s control and performance, which will be taken into consideration for future iterations on
the control system, displays, methods of control, and simulations. Data taken will also allow for quantitative
evaluation of ease of control and aircraft performance.
10
Future Plans
Planned follow-on technology maturation efforts include a large scale ground test program featuring full scale,
full propulsive flowpath wind tunnel testing with an actual engine, advanced structural design, and actuation to
measure loads, engine operability, thermal characteristics, and lift performance. Currently this effort is unfunded.
This program would achieve TRL 6 for the aero-propulsive configuration and allow for integration into a larger
platform acquisition effort in FY18.
Conclusions
The entire Speed Agile Concept Demo program was very successful, developing and validating integrated lift
systems, efficient transonic cruise technologies, and robust flight controls for a future theater airlifter with enhanced
cargo capacity and survivability. The hybrid powered lift systems have demonstrated superior lift performance to
enable access to short austere fields and delivery of relevant payloads closer to the point of need. The efficient
transonic cruise characteristics validated under this program have proven that faster, more aerodynamically efficient
transports of the future can save fuel, increase productivity, and quicken the response on time critical missions. The
development and evaluation of robust flight control schemes for STOL aircraft will allow safer, more precise
maneuvering and better field performance for highly coupled aero-propulsive vehicle concepts.
Acknowledgments
The accomplishments of the government and industry teams on this project were made possible by excellent
support from test centers at NASA (14’x22’, NTF, VMS), AEDC (NFAC), and AFRL (LAMARS). Williams
International, Advanced Technologies Incorporated, and Trimodels are thanked for their support in designing and
fabricating superior wind tunnel models and supporting equipment.
References
1
Zeune, C. H. “Enabling Speed Agility for the Air Force,” AIAA-2010-349, January 2010
Harrison, N. A., et. al. “The Design and Test of a Swept Wing Upper Surface Blowing Concept,” AIAA-2013-1102, January
2013
3
Shweyk, K. M. “Overview of Aerodynamic Model and Flight Control System of a Speed Agile Concept Demonstrator,” AIAA2013-1101, January 2013
4
Hooker, J. R., et. al. “Design and Transonic Wind Tunnel Testing of a Cruise Efficient STOL Military Transport,” AIAA-20131100, January 2013
5
Wick, A. T., et. al. “Powered Lift CFD Predictions of a Transonic Cruising STOL Military Transport,” AIAA-2013-1098,
January 2013
6
Barberie, F. J., et. al. “Low Speed Powered Lift Testing of a Transonic Cruise Efficient STOL Military Transport,” AIAA-20131099, January 2013
2
11

An Overview of the Air Force`s Speed Agile Concept Demonstration

Transcript

Documenti analoghi

The new high speed BOLOGNA - FLORENCE line

USB Adapter - Atlantis-Land

Mr. Ezio Facchin

Equipaggiamento Moto VFR 750 F (90 > 93)

Movotec Lift Systems

Lavoro con la Disney Cruise Line

kitesegnana kitesegnana

SENECA | Supporto e Assistenza

Diapositive 1

Europass CV - Felice Pescatore

HPR: water jet force measurement

1. Manual for the emulator EMUFDD

MDC DAC 2.0 USB Digital-to-Analogue Converter

seda 2015 book of abstracts

Iliac crest free flap in oromandibular reconstruction. 13 cases study