What is the purpose of the QSRA?

The QSRA is designed as a flight facility for advanced flight experiments in terminal area operations, specifically to gather data for the development of advanced STOL aircraft.

How does NASA minimize costs for research aircraft like the QSRA?

NASA minimizes costs by using existing airframes, off-the-shelf hardware, and applying cost-conscious management practices throughout the project.

When did the QSRA make its first flight?

The QSRA made its first flight on July 6, 1978, and the initial flight-test program was completed one month ahead of schedule.

What is the significance of the QSRA project for the U.S. economy?

The QSRA project is significant as it supports the U.S. aircraft industry's ability to export aircraft, which positively influences the U.S. balance of payments.

An overview of the quiet short-haul research aircraft program

Q: What will the data from the QSRA be used for?

The data will be used by the U.S. aircraft industry to establish design criteria and by regulatory agencies to set certification criteria for advanced STOL aircraft.

Document

NASA Technical Memorandum 78545 t [NASA-TM-78545) AN OVERVIEW OF THE Q U I E T W79- 11034 SHORT-HAUL RESEARCH AIRCRAFT PBOGBAI! (NASA) 41 p HC A03/flIF A01 CSCL 01c Unclas e3105 37159

An Overview of the Quiet Short-Haul

Research Aircraft Program

Michael D. Shovlin and John A. Cochrane

November 1978 National Aeronautics and Space Administration A N OVERVIEW OF THE QUIET SHORT-HAUL RESEARCH AIRCRAFT PROGRAM Michael D. Shovlin and John A. Cochrane Ames Research Center INTRODUCTION The Quiet Short-Haul Research Aircraft (QSRA) is a new research aircraft which NASA will use as a flight facility for advanced flight experiments in Because the nature and use of research aircraft by terminal area operations.

NASA are frequently misunderstood, the following discussion is presented to clarify the subsequent description of the QSRA and its use as a research facility .

NASA research aircraft are not prototypes and frequently they are not experimental aircraft. For example, the Kuiper Airborne Observatory, which is operated by Ames Research Center, is a standard Lockheed/USAF C-141 modi- fied to carry a telescope and other airborne scientific experiments. Other research aircraft such as the X-15 series are highly experimental in nature, but are not prototypes for future aircraft. Occasionally, an aircraft built as a prototype is used as a research aircraft; examples are the Boeing "Dash Eighty," which was the707prototype and the USAF AMST prototypes. These air- craft were used or planned to be used by NASA for flight research after com- pleting their prototype missions. All of these aircraft had a common denominator: as research aircraft their mission was one of data gathering, and indeed, this is usually the primary mission of NASA research aircraft.

For this reason, NASA frequently views research aircraft as facilities, just as a wind tunnel or a simulator is considered a test facility.

The data resulting from the QSRA flight research program will be used by the United States aircraft industry to establish design criteria and by regu- latory agencies to establish certification criteria for advanced STOL air- craft. This is important from a national point of view since aircraft exports exert an important positive influence on the U.S. balance of payments. In addition, QSRA flight data will lead to improved air transportation at reduced noise levels and with less air traffic congestion.

Another characteristics of many NASA research aircraft is lower cost than that typically associated with a prototype development. Limited budgets and fiscal responsibility dictate that research capability must be maximized rela- tive to cost, and experience with research airplane projects has led to cer- tain approaches developed to minimize their cost. These include: 1 . Use of an existing airframe where possible 2. Use of "off-the-shelf" hardware 3. Use of goals instead of requirements 4.

In-house participation where appropriate 5.

Cost consciousness at all organizational levels 6 .

Soft tooling 7 . Informal documents The application of these concepts to the QSRA will be discussed later.

HISTORICAL BACKGROUND NASA has conducted research with powered-lift airplanes since the 1950s.

The first jet STOL research airplane developed by NASA was the Augmented Jet Flap STOL Research Airplane (ref. 1) developed in the early 1970s. This was an austere program which consisted of a modification of an existing deHavilland C-8A Buffalo, powered by two modified Rolls Royce Spey engines.

It recently completed 500 hr of highly successful flight research and after a major inspection has been placed back in service for futher work. Its performance is representative of the first generation of jet STOL aircraft with an approach lift coefficient in the 3.5 to 4.0 range. Its major limita- tion is high levels of sideline noise.

A second, more ambitious jet STOL research airplane program, initiated in the early 1970s was the Quiet Experimental STOL airplane (QUESTOL). Three preliminary design studies and a design competition (won by Lockheed-Georgia) were completed before this project was cancelled in January 1973 due to budgetary limitations. The QUESTOL was planned as a four-engine, externally blown-flap STOL airplane, powered by four General Electric TF-34 turbofan engines.

In January 1974, a decision was made to embark on an austere jet STOL research aircraft which would feature very low sideline noise levels and "next generation'' performance (approach lift coefficient of 4.5 to 5 . 5 ) . Prelimin- ary design contracts were awarded the Boeing Commercial Airplane Company and the Lockheed-Georgia Company to study an augmented jet-flap concept and a hybrid upper surface blowing concept. These studies were not competitive in the sense that they did not form the basis for the selection for the winner of the subsequent hardware competition. Each design team operated indepen- dently, and only at an industry-wide conference at the end of the study was the work of one team revealed to the other. In this way, NASA was able to obtain two independent approaches to the problem. The results of these studies were presented to industry in September 1974 (refs. 2, 3 ) , and a QSRA was request for proposal for detail design, fabrication, and test of the issued in November 1974. Important excerpts from the initial statement of work are given in table 1. Boeing, Douglas, and Lockheed responded to the request for proposal and after a lengthy evaluation, the Boeing Commercial Airplane Company was awarded the hardware contract in March 1976; The QSRA made its first flight on schedule-July 6, 1978. The aircraft departed Boeing Field in Seattle to go to Paine Field in Everett, Washington Although the in order to begin its initial 17.5-hr flight-test program.

primary objective of this program was to demonstrate the airworthiness of the aircraft and its systems, the last 2 to 2.5 hr were devoted to internal and external noise measurements. This initial flight test program proceeded very smoothly and was accomplished 1 month ahead of schedule, allowing the aircraft to be delivered in August instead of in September of 1978 as originally planned. Figure 1 shows the aircraft on final approach into Moffett Field, California where 9.5 hr were flown in order to verify data system operation at Ames and to provide pilot familiarization and training.

After a thorough inspection of the aircraft and its engines, the second phase of the NASA flight research program was begun in November at Ames where the propulsive- lift, handling, and acoustic characteristics are being investigated, with imrxovement modifications as required.

MANAGEMENT APPROACH The total funding for the QSRA was established at $29 million in January 1974. At the start of the project, a firm commitment was made to complete Since these funds had to cover costs the project within the available funds.

for all studies, test, engine and airframe procurement, and proof-of-concept flight tests, an austere and innovative management approach was required.

The approximate distribution of the available money is given in table 2.

Scope Versus Cost The techniques discussed in the Introduction were all applied to the task of developing a technically meaningful project within the budgetary limita- tions. An important aspect of living within the budget was definition of the scope of the project, which was largely accomplished during the preliminary design studies. Within limits, a project such as QSRA can "cost what you want it to cost." This is due to the fact that many features, while highly desirable, are not essential. An example of this occurred during the prelim- inary design studies. An article in a trade magazine indicated that $32 million were available for the QSRA project. The first cost estimates, inde- pendently prepared by the two study contractors, were for about $30 million.

A special trip was made to NASA project managers to explain the distribution of funds as shown in table 2. The second round cost estimates were about $20 million. Both estimates were legitimate; the difference was in the scope and detail of the tasks to be accomplished.

There is, however, a lower limit beyond which a technically meaningful cost cannot be implemented. Some fortunate circumstances contributed to the completion of the QSRA project within the available funds. These included: ( 1 ) the availability of a suitable airplane, the deHavilland C-8A, for modifi- cation; and ( 2 ) the availability of suitable engines which could be configured for use in the QSRA.

Airframe Acquisition The C-8A w a s obtained at no c o s t , through a p p r o p r i a t e government chan- n e l s , from t h e National Center f o r Atmospheric Research (NCAR).

The p h y s i c a l s i z e and t h e T - t a i l c o n f i g u r a t i o n of t h e 6-8A made it an i d e a l airframe f o r modification i n t o an advanced STOL a i r c r a f t and p r i o r experience w i t h a s i m i - lar modification, t h e Augmented Jet Flap STOL Research A i r c r a f t (AWJSRA), f u r t h e r enhanced its d e s i r a b i l i t y .

When t h e F a i r c h i l d A-10 a i r p l a n e w a s s e l e c t e d as t h e winner of t h e United States A i r Force AX fly-off competition, t h e two Northrop A-9A a i r p l a n e s were t r a n s f e r r e d t o N A S A f o r a p o s s i b l e f l i g h t research program. A later d e c i s i o n not t o f l y t h e two prototype a i r p l a n e s made t h e engines, equipment, and s p a r e s from t h i s program a v a i l a b l e f o r QSRA use. S i x Lycoming YF-102 engines and four accessory power packages w e r e salvaged from t h e A-9A program together with many o t h e r miscellaneous components. The YF-102 engines, although r e l a - t i v e l y immature prototype engines, w e r e almost i d e a l f o r t h e QSRA. They are high by-pass r a t i o , geared-fan engines t h a t generate 33,410 N (7,500 l b ) of t h r u s t at low n o i s e l e v e l s .

The use of t h e C-8A and t h e YF-102 engines w a s an important f i r s t s t e p i n minimizing t h e c o s t of t h e QSRA p r o j e c t , b u t many a d d i t i o n a l cost-reduction f a c t o r s w e r e necessary. They included: in-house p a r t i c i p a t i o n , cooperative approach, d e t a i l e d t r a c k i n g of c o s t s , and a f u l l a p p r e c i a t i o n of t h e impor- t a n c e of c o s t a t a l l o r g a n i z a t i o n a l l e v e l s .

In-House P a r t i c i p a t i o n p a r t i c i p a t i o n w a s t o l e t N A S A do t h a t which The approach t o in-house N A S A could do b e s t and t o l e t Boeing do those t h i n g s which Boeing could do b e s t . The large-scale wind-tunnel model i s an e x c e l l e n t example of t h i s con- cept. Early i n t h e program, a large-scale, powered wind-tunnel model w a s i d e n t i f i e d as a p r o j e c t requirement. Ames Research Center h a s a long h i s t o r y of c o n s t r u c t i n g l a r g e powered models f o r research i n t h e A m e s 40- by 80-Foot Wind Tunnel. Boeing, on t h e o t h e r hand, had a d e t a i l e d knowledge of t h e d e t a i l s of t h e QSRA design. I n order t o t a k e advantage of t h e e x p e r t i s e of each o r g a n i z a t i o n , Boeing w a s assigned t h e t a s k of designing t h e model and N A S A assumed r e s p o n s i b i l i t y f o r f a b r i c a t i n g and instrumenting t h e model.

t h e engine program. An Another example of a N A S A in-house program w a s completed by t h e L e w i s Research Center i n e x t e n s i v e ground test program w a s of t h e QSRA which both performance and a c o u s t i c d a t a w e r e acquired i n support L e w i s a l s o managed t h e program t o r e f u r b i s h and update t h e design e f f o r t .

f l i g h t engines.

Cost Consciousness It i s beyond t h e scope of t h i s paper t o d i s c u s s t h e d e t a i l s of QSRA man- However, no agement beyond t h e examples t h a t have a l r e a d y been presented.

d i s c u s s i o n of QSRA management would b e complete without emphasizing t h e importance of t h e outstanding cooperation between t h e Boeing p r o j e c t team and t h e NASA P r o j e c t O f f i c e and t h e d e t a i l e d t r a c k i n g of c o s t s accomplished by both t h e s e groups. The Boeing p r o j e c t c o n t r o l group tracked c o s t f o r 38 work breakdown s t r u c t u r e elements on a weekly b a s i s a t t h e peak of t h e p r o j e c t .

These d a t a were provided i n a timely manner t o the NASA P r o j e c t O f f i c e and w e r e on d i s p l a y i n a c o n t r o l room i n t h e Boeing p r o j e c t area. Thus, p r o j e c t personnel t o t h e lowest o r g a n i z a t i o n a l levels were made aware of c o s t perfor- mance. The NASA P r o j e c t O f f i c e w a s consulted whenever t r a n s f e r s w e r e made from t h e Boeing management reserve. I n t h i s way, p r o j e c t funds were not expended t o c o r r e c t minor performance d e f i c i e n c i e s t h a t w e r e not important t o and a v a i l a b l e resources could be concentrated on important problems. .

NASA, A paper planned f o r later p u b l i c a t i o n by t h e QSRA p r o j e c t personnel w i l l d e a l with t h i s s u b j e c t i n depth.

IN-HOUSE PROGRAMS Wind-Tunnel Tests The need f o r a c c u r a t e , large-scale, wind-tunnel t e s t i n g of t h e s p e c i f i c powered-lift a i r p l a n e c o n f i g u r a t i o n had been i d e n t i f i e d i n s t u d i e s previous t o t h e QSRA c o n t r a c t award ( r e f . 4 ) . I n o r d e r t o support t h e QSRA design e f f o r t , and t o reduce c o s t s and r i s k by u t i l i z i n g NASA t a l e n t and f a c i l i t i e s , an e x i s t i n g large-scale, wind-tunnel model w a s modified t o be aerodynamically similar t o t h e QSRA and t e s t e d i n t h e A m e s 40- by 8O-Foot Wind Tunnel. T h i s 0.55-scale model w a s used t o provide aerodynamic and loads d a t a f o r t h e QSRA design, and t h e c o n t r o l s t a b i l i t y d a t a f o r t h e f l i g h t simulation program.

Wind-tunnel model-The QSRA wind-tunnel model is shown mounted i n t h e Ames 40- by 8O-Foot Wind Tunnel, i n f i g u r e 2. T h i s model is powered by f i v e JT-15D turbofan engines. Four of t h e s e engines are mounted above t h e wing; t h e f i f t h , mounted i n t h e f u s e l a g e , provides boundary-layer c o n t r o l (BLC) air.

The model has t h r e e trailing-edge f l a p systems. Upper-surface-blown f l a p s are l o c a t e d d i r e c t l y behind t h e engines, w i t h double-slotted f l a p s outboard of t h e s e and blown a i l e r o n s at t h e wing t i p s . The e n t i r e leading edge i s blown f o r boundary-layer c o n t r o l , including t h e area between t h e n a c e l l e and fuse- lage. Although t h e leading edge f l a p s were f i x e d , t h e t r a i l i n g - e d g e f l a p systems and s p o i l e r s could b e remotely a c t u a t e d during t h e test runs.

This model had over 600 p r e s s u r e and temperature measuring p o i n t s i n Engine order t o provide a i r l o a d s and temperature design d a t a f o r t h e QSRA.

t h r u s t levels w e r e measured ynder s t a t i c conditions w i t h f l a p s up and corre- l a t e d w i t h f a n speed. These c o r r e l a t i o n equations were used t o determine t h r u s t levels during t h e wind-tunnel test p o i n t s . The c o r r e c t e d mass flows w e r e obtained from i d e a l m a s s flows, c a l c u l a t e d by using t h e average s t a t i c p r e s s u r e a t t h e f a r t h e s t downstream i n l e t measuring p o i n t and t h e test s e c t i o n t o t a l temperature and pressure. This i d e a l flow w a s c o r r e c t e d f o r p r e s s u r e recovery and i n l e t l o s s e s by assuming a 0.98 c o r r e c t i o n f a c t o r .

Aircraft design contributions-The wind-tunnel tests of this model made a number of important contributions to the final QSRA design.

These tests defined the airload data used in the QSRA structural design, showing that these loads were slightly different in location and magnitude than those pre- dicted using YC-14 data, particularly on the fuselage near the wing leading edge. These tests also verified the mixed flow nozzle and propulsion system simulation which was based on Langley JT-15D tests simulating the YC-14 pro- pulsion system. The tests defined the BLC system requirements, showing that blowing was necessary at all times along the leading edge during high angle- of-attack operations, resulting in a redesign of the QSRA's BLC system.

These tests defined and verified the effectiveness of the control surfaces and the effects of engine-out and other failure conditions, providing a data base for the flight simulation. Finally, these tests defined a serious nacelle/wing aerodynamic interference problem and provided a simple, effective, low-cost solution for the aircraft design, by showing that several small vortex gener- ators could eliminate the boundary-layer separation at this interface. A more detailed summary of the test data is contained in references 5-7.

Configuration optimization-A continuing benefit of these wind-tunnel tests is in the use of this data base and the model as a tool for further con- figuration development and optimization. As an example, the early wind-tunnel tests showed that although blowing was essential along the leading edge at high angles-of-attack, only very small amounts were required to keep the flow attached over the wing. In subsequent tests, a slotted leading-edge flap was fabricated and tested to determine what performance penalties, if any, were associated with removing the outboard leading edge BLC system (ref. 5 ) .

4 ' in Although performance improved slightly, there was a loss of about angle-of-attack margin. One of the projected studies to be made with the QSRA will be to verify in flight the effect of replacing the outboard leading edge BLC system,witha slotted flap, a change that would considerably simplify the aircraft pneumatic system. This change will be made, however, only after a thorough documentation of the flying characteristics of the basic configura- tion.

Engine Ground Tests As discussed previously, the QSRA is powered by the Lycoming YF-102 engines acquired from the A-9A aircraft program. Although these are rela- tively immature prototype engines, they had met all of their performance goals during the AX program, and had demonstrated operational reliability. The QSRA, however, required a much more complex engine installation with a confluent 10% of core flow exhaust system and with a bleed air schedule requiring up to QSRA requirements were so far beyond airflow at low power settings. These the existing engine performance data base that there were questions regarding In addition, the engine operation and its effect on cost and aircraft safety.

ambitious acoustic goals of the QSRA required an extensive acoustic data base in order to develop an adequate low-noise nacelle design within cost con- straints. It became obvious that it was necessary to develop these data bases in order to minimize program cost and risk, and again it was clearly an area where NASA talent and facilities could be used most effectively.

Vertical L i f t Fan Facility-The L e w i s Research Center Vertical L i f t Fan F a c i l i t y is an outdoor engine test s t a n d s h e l t e r e d by a service b u i l d i n g which The engine is suspended beneath t h e is moved away on t r a c k s b e f o r e t e s t i n g .

t h r u s t measuring system, which can b e pivoted around a vertical a x i s f o r oper- a t i o n a l f l e x i b i l i t y . A frame work extending from t h e t h r u s t measuring system is used t o mount i n l e t and exhaust hardware s e p a r a t e l y from t h e engine.

The engine c e n t e r l i n e was 2.9 m (9.5 f t ) above t h e ground; t h e f a c i l i t y , with t h e b a s e l i n e confluent flow YF-102 mounted on t h e t h r u s t s t a n d , is shown i n f i g u r e 3. The area beneath t h e engine is paved w i t h concrete and a s p h a l t o u t t o t h e a c o u s t i c d a t a microphones which are l o c a t e d on a 30-m (100 f t ) r a d i u s circle over a 1 6 0 ' a r c from t h e i n l e t c e n t e r l i n e . The c o n t r o l room is l o c a t e d about ' 152 m (500 f t ) from t h e stand and a f f o r d s a good view of t h e engine i n l e t .

ProDulsion design refinements-The engine ground test c o n t r i b u t e d t o t h e f i n a l QSRA a i r c r a f t design i n a number of d i f f e r e n t ways. The tests were used t o d e f i n e t h e engine base-line performance f o r t h e confluent flow config- u r a t i o n and t o update t h e engine performance p r e d i c t i o n deck. These tests defined t h e t r a n s i e n t o p e r a t i n g c h a r a c t e r i s t i c s of t h e engine and t h e e f f e c t of BLC system f a i l u r e s on t h e engine s t a b i l i t y and s a f e operation. These tests showed t h a t t h e engine w a s unable t o a c c e l e r a t e from low power s e t t i n g s under high c o r e bleed conditions, r e q u i r i n g t h e design of a BLC c o n t r o l system t h a t l i m i t s core bleed t o power s e t t i n g s where t h e engine can be s a f e l y oper- a t e d . New acceptance test procedures w e r e developed as a r e s u l t of t h e s e ground tests i n order t o ensure adequate, s t a b l e , and s a f e engine operation when i n s t a l l e d i n t h e QSRA. F i n a l l y , a s p e c i a l test w a s run t o v e r i f y t h e design and adequate operation of t h e fan bleed air S-duct and i t s flow char- a c t e r i s t i c s a t t h e e j e c t o r i n l e t ( r e f . 8 ) .

Acoustic design refinements-Acoustic performance is a second area where t h e L e w i s test program made s i g n i f i c a n t c o n t r i b u t i o n s t o t h e QSRA design.

These tests developed a l l of t h e a c o u s t i c d a t a base f o r t h e YF-102 engine, providing a measure of t h e component n o i s e l e v e l s and t h e i r d i r e c t i v i t y . The induct f a n tones and t h e i r l o c a t i o n r e l a t i v e t o t h e duct w a l l s w e r e determined along with t h e i r mode shapes and o t h e r design d a t a . Acoustic design simpli- f i c a t i o n s e l i m i n a t i n g s p l i t t e r r i n g s and engine spinner treatment w e r e veri- f i e d , e l i m i n a t i n g considerable c o s t and performance p e n a l t i e s . The e f f e c t of wing s h i e l d i n g w a s determined and, f i n a l l y , s p e c i a l techniques w e r e developed t o determine t h e c o n t r i b u t i o n of combustor n o i s e ( r e f . 9).

F l i g h t Simulation The A m e s F l i g h t Simulator f o r Advanced A i r c r a f t (FSAA) w a s used t o d e f i n e t h o s e combinations of f l i g h t conditions, a i r c r a f t c o n f i g u r a t i o n , con- rates t h a t would ensure a c c e p t a b l e handling q u a l i t i e s t r o l power, and c o n t r o l f o r both normal operation and i n v a r i o u s s i n g l e o r m u l t i p l e f a i l u r e occur- rences i n e i t h e r propulsion o r f l i g h t c o n t r o l systems.

FSAA-The FSAA is a six-degree-of-freedom motion s i m u l a t o r w i t h very high f i d e l i t y motion and v i s u a l cues. It has two p i l o t s t a t i o n s and room f o r a n It w a s configured t o c l o s e l y approximate t h e f l i g h t observer i n t h e cab.

deck of the QSRA with similar instruments, throttles and controls.

A pilot's 4 .

eye view of the simulation is shown in figure Experience with the FSAA has shown that realistic, accurate simulations can be made and, indeed, the 4-5 for a single pilot ratings of 2-3 for normal operation at low speed and failure were verified in flight.

These simulations showed a need for several design changes to improve handling qualities under a variety of STOL operations and simulated failure conditions. The need for longitudinal stability augmen- tation and direct lift control was identified, as was a change in horizontal stabilizer incidence. A requirement was also determined for automatic retrac- tion of the upper surface blown flaps to reduce drag during go-around. Pilot procedures and handling qualities were also defined for operations with one or more engines inoperative, and for situations where electrical power was lost, or hydraulic or boundary-layer control systems had failed. Steep curvi- linear landing approach operating procedures were investigated for noise abatement .

Further details of the QSRA flight simulations, the QSRA mathematical model, and the results of these simulations are contained in references 10-12.

AIRPLANE DESCRIPTION The general arrangement of the QSRA is shown in figure 5 and a photograph of the airplane is shown in figure 1. The fuselage is that of a deHavilland C8-A Buffalo with structural reinforcement in the aft fuselage and new fair- ings at the wing-body intersection. The C-8A empennage was used without struc- tural or aerodynamic modification. SAS actuators were added to both the rudder and the elevator and a hydraulic actuator was added for power actuation of the elevator. The C-8A landing gear was modified to increase the sink rate capa- bility of the aircraft.

The QSRA wing was designed and fabricated bz Boeing with a wingspan of

22 .4 m (73.5 ft), a wing area of 55.74 m2 (600 ft >, and a quarter chord sweep

of 15O. Figure 6 shows the wing being attached to the fuselage at the Boeing Development Center in Seattle. The center section of the wing is sealed to form two integral fuel cells which contain a total of 4535.9 kg (10,000 lb) of Jet A-1 (JP-5) fuel. Fixed leading edge flaps are blown by a mixed flow boundary layer control system. The trailing edge on either side of the cen- terline consists of two upper surface blowing (USB) flaps, a double-slotted flap, and a drooped, blown aileron.

In The flaps and ailerons are supported by external beams and linkages.

keeping with the austere nature of the program and the low-speed environment The main landing gear is fixed and is of the QSRA, these are not faired.

attached to the underside of the wing between the two nacelles. The wing is attached to the fuselage by the same pin joints as those used in the original C-8A. This provided a significant cost saving but it did require the addition of 4 1 8 . 7 kg (923 lb) of ballast in the tail.

Propulsion System The QSRA main propulsion system consists of four AVCO-Lycoming YF-102 (QSRA) engines mounted in above-the-wing nacelles (fig. 5 ) . These prototype engines, acquired from the A-9A program, were extensively refurbished and updated in a program managed by the Lewis Research Center. The principal elements of this update include a fan containment ring, combustor case high- pressure air bleed ports, new oil coolers, and improved shafting material.

Powerplant-A cutaway view of the engine is shown in figure 7. The low- pressure spool incorporates a single-stage fan which provides bypass and core, air to the engine. The core airflow is further compressed by a single-stage supercharger attached to the fan. The fan is driven by a two-stage, uncooled turbine through a single planetary reduction gear (2.3 speed ratio) located in the fan module.

The gas producer section of the engine is essentially a T-55 core with slight modifications. The high-pressure components include an axial/centrifu- gal compressor, a reverse-flow combustor, and a two-stage, air-cooled turbine to drive the compressor. The high-pressure compressor has seven axial stages followed by a centrifugal stage. It features variable inlet guide vanes (VIGV'S) and a sixth-stage bleed band to minimize the possibility of compres- sor stall during transient operations.

The engine weighs 5412 N (1215 lb) and has a basic diameter of 1.077 m (42.4 in.) with an overall length of 1.621 m ( 6 3 . 8 in.) including the fan spinner, while the fan has a diameter of 1.024 m ( 4 0 . 3 i n . ) . The engine geom- etry and unstalled performance are shown in figure 8 .

Nacelle structure-The nacelle layout is shown in figure 9 and the major external nacelle structural components are shown in figure 10. The external nacelle is composed of two main assemblies, the structural cowl and nozzle assembly and the engine build-up assembly. The structural cowl and nozzle assembly is attached to the wing front spar, forming the structural nacelle The engine build-up assembly is then mounted to this structure and pylon.

and forms the front half of the nacelle. The nose cowl is attached to the engine as shown in figure 1 1 and forms the inlet and outer nacelle. In addi- tion to the nose cowl, a core cowl and the primary nozzle are installed as part of the engine buildup. Engine driven accessories are airframe-mounted in the nose cowl, resulting in this area being one of the three primary fire zones in the nacelle. Other fire zones are the core cavity and the outer fan case. Fire protection behind the nacelle is provided by an external heat shield attached to the upper surface of the wing, together with the use of In addition, heat-resistant materials in the wing flaps and trailing edge.

the primary nozzle is canted bpwards allowing cooling fan air to be drawn between the wing surface and the high-temperature jet in normal operation.

Exhaust nozzle-The QSRA/YF-102 exhaust system is a confluent-flow design with both primary and fan streams discharging through a common D-shaped exit nozzle having an aspect ratio of 3.5. As indicated in figure 12, the core exhaust diffuses as it passes through the primary nozzle and then mixes with t h e surrounding f a n stream, e x i t i n g through t h e D-shaped upper-surface blowing nozzle. The core nozzle is canted upward 9.4' r e l a t i v e t o t h e engine center- l i n e t o minimize t h e h e a t e f f e c t s on t h e wing and f l a p s .

The flow areas i n t h e fan-duct and core-nozzle e x i t plane (mixing plane) are chosen t o provide adequate performance without s i g n i f i c a n t l y a f f e c t i n g The main c o n t r o l on s u r g e margins and engine match, however, surge margins.

is provided by t h e f i n a l e x i t area of t h e D-nozzle, which is designed t o spread t h e exhaust i n t o a t h i n s h e e t , which is then turned by t h e Coanda e f f e c t over t h e USB f l a p s , providing l i f t .

The QSRA D-duct design h a s mixing plane areas of 0.44 m2 (682.5 i n . 2 ) and 0.156 m2 (250 i n . 2 > f o r t h e f a n and core airstreams, r e s p e c t i v e l y . The D-nozzle w a s designed s o t h a t t h e e x i t area could be increased as much as l o % , if needed, from an e f f e c t i v e area of 0.42 m2 (650 i n . 2 ) which w a s 7.5% under t h e assumed b a s e l i n e area. Subsequent c a l c u l a t i o n s showed t h a t t h e r e would be adequate surge margins a t t h i s area, however, and t h a t no nozzle t r i m would be needed. Measured r e s u l t s discussed i n a later s e c t i o n showed t h a t indeed t h i s w a s t h e case.

Acoustic treatment-The l o c a t i o n of t h e n a c e l l e a c o u s t i c l i n e r s is shown i n f i g u r e 12. These l i n e r s are l o c a t e d i n two d i f f e r e n t n a c e l l e areas, t h e f a n duct and t h e i n l e t . The f a n duct l i n e r s , which are l o c a t e d on both t h e s t r u c t u r a l cowl and on t h e core cowl, are composed of p e r f o r a t e d aluminum f a c e s h e e t s bonded t o a n aluminum honeycomb core with s o l i d aluminum o u t e r backing s h e e t s . These panels cover about 0.75 m (30 i n . ) of d u c t l e n g t h and are estimated t o provide about 1 2 PNdB of a f t f a n a t t e n u a t i o n . These panels s e r v e as an i n t e g r a l p a r t of t h e cowl s t r u c t u r e and are load-carrying i n a d d i t i o n t o providing sound a t t e n u a t i o n .

The second area of t h e n a c e l l e t h a t i s l i n e d is t h e i n l e t . The i n l e t a c o u s t i c panels are double-layer c o n s t r u c t i o n with p e r f o r a t e d aluminum f a c e a s o l i d aluminum backing s h e e t and septum w i t h aluminum honeycomb c o r e s and s h e e t . The lower a c o u s t i c panel honeycomb c o r e s are s l o t t e d and d r a i n h o l e s water accumulation and p o s s i b l e are provided i n t h e o u t e r s h e e t t o prevent f r e e z e damage.

BLC System A unique f e a t u r e of t h e QSRA is t h e mixed-flow boundary-layer c o n t r o l A i r f o r t h e BLC system (BLC) system f o r t h e wing leading edges and a i l e r o n s .

A is b l e d from both t h e f a n and t h e engine c o r e and mixed i n an e j e c t o r .

schematic of t h i s system is shown i n f i g u r e 13.

The BLC a i r is d i s t r i b u t e d by cross-ducting from each engine t o t h e oppo- Interconnecting site s i d e of t h e wing leading edge o r a i l e r o n s u r f a c e s .

ducting and check valves are i n s t a l l e d between t h e two BLC system manifolds; t h e outboard n a c e l l e s .

they are l o c a t e d e x t e r n a l l y under t h e wing outboard of The a i l e r o n ducting is l o c a t e d i n a c a v i t y a f t of t h e rear s p a r ; however, t h e leading edge ducting had t o be l o c a t e d e x t e r n a l l y behind t h e leading edge A s w a s previously f l a p s and c r o s s e s over i n s i d e t h e f u s e l a g e , under t h e wing.

discussed, t h e s e ducts may be s i m p l i f i e d , i n a f u t u r e test period, t o e l i m i n a t e some of t h e e x t e r n a l ducts. The BLC ducting is arranged so t h a t each engine feeds a s e p a r a t e p a r t of t h e BLC system, as shown i n t a b l e 3 . This arrangement provides a degree of automatic r o l l compensation i n t h e event of t h e l o s s of a c r i t i c a l (outboard) engine.

Key elements i n t h e BLC system are t h e mixing ejector and servo-regulator The e j e c t o r , which is shown i n valve which are located as shown i n f i g u r e 1 4 .

t h e i n s e r t of f i g u r e 15, has a fixed-geometry mixing s e c t i o n with an e l l i p t i c a l c e n t e r body, and 42 c i r c u m f e r e n t i a l l y d i s t r i b u t e d e j e c t o r nozzles. These convergent-divergent nozzles, with length-to-diameter r a t i o s of 5:1, l i m i t .

t h e high-pressure bleed t o a nominal 10% of t h e engine core flow; and fan bleed i s l i m i t e d t o 3% due t o duct s i z e . Figure 15 shows t h e e f f e c t of t h i s e j e c t o r design on n e t blowing momentum of t h e a i l e r o n nozzles. The upper curve repre- s e n t s t h e performance of t h e e j e c t o r without any pressure regulation. The servo-regulator valve l i m i t s t h e downstream duct p r e s s u r e t o a preset value, however, and t h e regulated system follows t h e lower curve of f i g u r e 15, yield- ing a n e a r l y constant value of blowing momentum over t h e e n t i r e engine t h r u s t range. This valve r e g u l a t e s high-pressure flow from t h e compressor so t h a t i t is zero a t high power s e t t i n g s where t h e f a n pressure r a t i o is high, and about 10% of t h e c o r e a i r f l o w a t low power s e t t i n g s . Although t h e r e is a l o s s i n engine t h r u s t a t t h e 10% bleed a i r f l o w , it only occurs when a low-thrust l e v e l is commanded by the p i l o t . A t high-thrust s e t t i n g , t h e t h r u s t l o s s i s less than 1% due t o BLC system losses.

F l i g h t Controls The f l i g h t c o n t r o l s u r f a c e s are shown schematically i n f i g u r e 16. A l l wing t r a i l i n g edge s u r f a c e s are e l e c t r i c a l l y c o n t r o l l e d (fly-by-wire) except t h e a i l e r o n s . The s p o i l e r s , double-slotted f l a p s , and U S B f l a p s are elec- t r i c a l l y commanded and h y d r a u l i c a l l y powered; t h e a i l e r o n s are mechanically c o n t r o l l e d and h y d r a u l i c a l l y powered. Both t h e rudder and e l e v a t o r a r e C-8A components which are mechanically c o n t r o l l e d and both are h y d r a u l i c a l l y pow- ered.

Flap system-The USB f l a p s are deployed t o t h e 30° p o s i t i o n with t h e p i l o t s f l a p lever. A thumb switch located i n t h e t h r o t t l e handle f o r t h e No. 1 engine c o n t r o l s USB f l a p p o s i t i o n from 30° t o t h e f u l l d e f l e c t i o n of 66O. This provides t h e p i l o t with a convenient means of varying USB f l a p s e t t i n g , during a landing approach, as a means of g l i d e path control. Deploy- ment of t h e double-slotted f l a p s is c o n t r o l l e d by a s e p a r a t e l e v e r on t h e p i l o t ' s console. The a i l e r o n droop is slaved t o the double-slotted f l a p s .

The U S B f l a p s , t h e s p o i l e r s , and t h e double-slotted f l a p s are a l l i n d i v i d u a l l y actuated by d i g i t a l , e l e c t r o n i c a l l y c o n t r o l l e d , hydraulic a c t u a t o r s . This allows any f l a p o r s p o i l e r t o be actuated independently of any o t h e r by proper preprogramming. This f e a t u r e provides maximum research c a p a b i l i t y f o r t h e QSRA. A s i n i t i a l l y configured, t h e QSRA p i l o t h a s t h e c a p a b i l i t y t o command assymetric deployment of t h e double-slotted f l a p s t o t r i m engine-out r o l l i n g moment.

S t a b i l i t y augmentation systen-The QSRA has a s i n g l e channel, three-axis l i m i t e d a u t h o r i t y series type s t a b i l i t y augmentation system (SAS). The r o l l and yaw axes are s t a b i l i z e d by a s i m p l e analog system s i m i l a r t o t h e one used i n t h e Augmented Jet Flap STOL Research Airplane. The l o n g i t u d i n a l SAS is a rate-command, a t t i t u d e - h o l d system. It uses a General Electric MCP-701A d i g i - t a l computer t o provide both p i t c h SAS f u n c t i o n s and t o c o n t r o l t h e d i r e c t lift c o n t r o l system and c e r t a i n o t h e r l o g i c functions. When t h e d i r e c t l i f t c o n t r o l (DLC) f u n c t i o n is s e l e c t e d , t h e s p o i l e r s are deployed t o about -13'.

An i n c r e a s e i n t h r u s t ( t h r u s t levers forward) causes t h e s p o i l e r s t o retract, and a decrease i n t h r u s t causes them t o extend beyond t h e nominal -13' angle.

A f t e r each excursion from t h e -13' p o s i t i o n , a washout c i r c u i t gradually r e t u r n s t h e s p o i l e r s t o t h e -13' p o s i t i o n u n t i l t h e t h r o t t l e is once again moved. A t go-around t h r u s t levels, t h e DLC system is automatically disabled and t h e s p o i l e r s are r e t r a c t e d .

Additional information on t h e QSRA configuration and systems i s contained i n r e f e r e n c e 13.

AIRCRAFT PERFORMANCE Although a l a r g e number of ground and f l i g h t tests have been performed t o determine t h e QSRA's operating c h a r a c t e r i s t i c s , many of t h e s e d a t a are still being analyzed at t h i s w r i t i n g . Therefore, many of t h e a i r c r a f t performance curves presented are based on p r e d i c t e d , wind-tunnel, o r simulation r e s u l t s .

Most of t h e ground test d a t a have been analyzed, however, allowing a compari- son of t h e p r e d i c t e d and measured C h a r a c t e r i s t i c s of t h e propulsion and BLC systems. I n a d d i t i o n , s u f f i c i e n t f l i g h t d a t a have been checked t o v e r i f y t h a t t h e a i r p l a n e performance i s c l o s e t o t h a t p r e d i c t e d , and t h e s e f l i g h t r e s u l t s w i l l be commented on i n t h e p r e s e n t a t i o n of t h e i n d i v i d u a l d a t a curves.

Propulsion System Most of t h e measurements of propulsion system c h a r a c t e r i s t i c s w e r e made during t h e Boeing ground test. The primary o b j e c t i v e s of t h i s ground test (1) t o determine t h e component map c h a r a c t e r i s t i c s and t o v e r i f y ade- w e r e : q u a t e s u r g e margins (nozzle t r i m ) ; (2) t o measure engine performance with and without t h e BLC system o p e r a t i n g ; ( 3 ) t o t r i m and a d j u s t t h e engine f u e l con- t r o l s f o r i d l e , takeoff power, and acceptable a c c e l e r a t i o n c h a r a c t e r i s t i c s ; and, ( 4 ) t o measure flaps-up t h r u s t and flaps-down turning. I n order t o m e e t t h e s e o b j e c t i v e s , a l a r g e number of d a t a p o i n t s w e r e taken and analyzed f o r each engine; however, it w i l l only b e p o s s i b l e t o p r e s e n t a s m a l l representa- of t h e d a t a here. Four summary p l o t s w i l l be presented. Two t i v e sample d e a l with engine t h r u s t r e l a t i o n s h i p s , one w i t h t h e l o c a t i o n of t h e engine o p e r a t i n g l i n e s on t h e f a n map, and t h e f o u r t h with t h e a c c e l e r a t i o n charac- teristics of t h e engines with t h e BLC bleed schedule.

Thrust characteristics-The r e l a t i o n s h i p of t h e engine t h r u s t with fan speed is shown i n f i g u r e 1 7 , which a l s o shows t h e r e l a t i o n s h i p between fan and c o r e speed f o r t h e s e engines. This curve is based on t h e ground test r e s u l t s of a l l f o u r of t h e engines ( t h e r e s u l t s had less than tl% s c a t t e r ) .

The c o r r e l a t i o n between t h e p r e d i c t e d performance is very good above 72% of t h e c o r r e c t e d c o r e speed but very poor a t lower core speeds; hence, t h i s curve g i v e s b e t t e r r e s u l t s than computer deck and w i l l be used f o r performance esti- mation i n t h e f l i g h t test program. As measured i n t h e ground test, t h e in- board engines have about 3% less i n s t a l l e d t h r u s t than t h e outboard engines which have a maximum i n s t a l l e d t h r u s t of 30,068 N (6,750 l b ) .

The e f f e c t of ambient temperature a t sea level on t h i s maximum i n s t a l l e d t h r u s t is shown i n f i g u r e 18. These takeoff d a t a are shown f o r no-bleed and ECS-bleed only because t h e high p r e s s u r e bleed is normally s h u t o f f a t t h i s power s e t t i n g .

One a d d i t i o n a l i t e m t o n o t e is t h e r e s t r i c t e d zone shown i n f i g u r e 1 7 . T h i s , r e s t r i c t i o n r e s u l t s from a resonance problem i n t h e sun gear a t a f a n speed of 55.5%. Although t h i s resonance has a very sharp peak, t h e r e s t r i c t e d oper- a t i n g band h a s been set between 50 and 60% t o prevent excessive excursions i n t o t h i s zone, p a r t i c u l a r l y near t h e resonance peak. This band causes a b a s i c problem i n t h e a i r c r a f t operation by f o r c i n g t h e STOL f l i g h t i d l e set- t i n g t o be a t a nominal 60% which i s about 5-6% higher than is d e s i r a b l e .

AVCO-Lycoming is working on a redesign of t h e sun gear t o e l i m i n a t e t h i s resonance problem and expects t o have a s o l u t i o n some t i m e i n 1979.

Fan operation-Figure 19 shows test d a t a from a l l four engines p l o t t e d on t h e YF-102 f a n map. This map includes p r e d i c t e d operation f o r a range of areas with t h e base area corresponding t o t h e untrimmed noz- r e l a t i v e nozzle z l e s .

The test d a t a w e r e p l o t t e d using f a n p r e s s u r e r a t i o and corrected bypass flow as primary parameters with c o r r e c t e d fan speed as a secondary parameter. The test d a t a i n d i c a t e d good c o r r e l a t i o n w i t h p r e d i c t i o n s up t o f a n speeds of 70%; a t h i g h e r speeds, t h e s e d a t a i n d i c a t e d a nozzle under-area condition of up t o 2%. This nozzle area spread w a s considered good i n l i g h t of t h e r a t h e r l i m i t e d instrumentation and hence a d e c i s i o n w a s made t o oper- ate without a d d i t i o n a l nozzle t r i m . The upper l i m i t f o r fan operation used during t h e A-9A program's YF-102 acceptance tests is a l s o shown i n t h e f i g u r e f o r reference.

Engine acceleration-Engine a c c e l e r a t i o n s w e r e i n i t i a l l y conducted on a l l engines s t a r t i n g a t t h r e e f a n speeds-48, 53, and 60%-which encompassed t h e p r e d i c t e d STOL f l i g h t i d l e range. Adjustments w e r e made, t o t h e f u e l c o n t r o l s of a l l t h e engines, t o i n c r e a s e t h e a c c e l e r a t i o n schedule i n order t o i n c r e a s e t h e a c c e l e r a t i o n rate w i t h t h e Boeing high-pressure bleed schedule. Figure 20 shows t h e t h r u s t v e r s u s t i m e f o r a c c e l e r a t i o n s from 53% f a n speed f o r a l l t h e engines w i t h t h e f i n a l f u e l c o n t r o l adjustment. A l l of t h e engines a c c e l e r a t e t o 65% t h r u s t i n approximately t h e same t i m e ; however, from that p o i n t on engine N o . 2 w a s markedly slower and took s e v e r a l more seconds than t h e o t h e r engines t o come up t o t h e 95% t h r u s t p o i n t . This slowness is due t o a n i n t e r - valve on engine No. 2 begins t o c l o s e a t a c t i o n w i t h t h e BLC system, where t h e a higher speed and appears t d c l o s e at a slower rate. This r e s u l t s i n higher bleed rates at a given f a n speed f o r engine No. 2 which reduces t h e accelera- t i o n rate a t higher f a n speeds. Engine No. 1 a l s o e x h i b i t s t h i s c h a r a c t e r i s - t i c but t o a much lesser e x t e n t than engine No. 2.

I n i t i a l l y , t h e s e a c c l e r a t i o n d a t a w e r e t o be used t o e s t a b l i s h t h e STOL However, due i d l e d e t e n t p o s i t i o n which corresponded t o a fan speed of 53%.

t o t h e problems discussed i n t h e previous s e c t i o n , t h e t h r o t t l e has no f l i g h t i d l e d e t e n t and normal STOL landing operation is a t 60% of fan speed. Because of t h e geometry of t h e overhead t h r o t t l e system, t h e 60% s e t t i n g forms a "nat- u r a l " f l i g h t i d l e p o s i t i o n .

Engine s t a b i l i t y - S t a b i l i t y tests w e r e conducted on a l l engines t o demon- strate acceptable i n l e t o p e r a t i o n and acceptable surge margins w i t h increased f u e l c o n t r o l a c c e l e r a t i o n schedules. These tests w e r e conducted with a l l bleeds off and t h e f u e l flow increased by 5% (TEST p o s i t i o n f o r compressor surge d e t e c t i o n ) . To check s t a b i l i t y under s e v e r e o p e r a t i n g conditions a series of t r a n s i e n t s , c o n s i s t i n g of r a p i d a c c e l e r a t i o n s , d e c e l e r a t i o n s , and Bodies, were performed on t h e engines. No adverse engine o p e r a t i o n w a s noted .

and t h e engines operated surge-free during t h i s e n t i r e series of tests.

One f i n a l s t a b i l i t y test w a s conducted on t h e engine i n order t o check A wind machine w a s posi- i n l e t and f a n operation at high angles of a t t a c k .

tioned t o provide a 36 m / s (72 knot) wind at 51' t o t h e i n l e t c e n t e r l i n e which w a s estimated t o be t h e most c r i t i c a l i n l e t inflow condition. Tests w e r e run on engines Nos. 1 and 2 using t h e same procedures as i n t h e previous tests and a l s o with t h e engines a t ground i d l e (low m a s s flow) i n a crosswind, which i s t h e worst condition f o r i n l e t s e p a r a t i o n . Again no adverse engine operation w a s detected.

BLC System Performance An e v a l u a t i o n of t h e ground-test d a t a showed t h a t t h e BLC system perfor- mance w a s e s s e n t i a l l y as p r e d i c t e d w i t h t h e amount of n e t blowing momentum b e t t e r than o r equal t o p r e d i c t e d l e v e l s a t both STOL i d l e and takeoff power.

The operation of t h e high-pressure r e g u l a t o r valve w a s s t a b l e w i t h t h e pumping performance of each e j e c t o r compatible with i t s system demands. The perfor- w a s i n e x c e l l e n t agreement w i t h c a l c u l a t e d per- mance of t h e a i l e r o n system formance, both with and without t h e r e g u l a t o r valve working ( f i g . 15). The only d e v i a t i o n from t h i s curve occurred a t t h r u s t s e t t i n g s above 70% where system performance w a s s l i g h t l y higher than t h a t predicted. The test r e s u l t s showed t h a t system l o s s e s a t t h e design p o i n t were i n good agreement with pre- d i c t i o n s , t h e l o s s e s being 5.2 and 6.5 % of t h e mixing t o t a l p r e s s u r e f o r t h e leading edge and a i l e r o n systems, r e s p e c t i v e l y .

During t h e BLC system tests, t h e only s i g n i f i c a n t problem t h a t w a s encountered w a s t h e i n a b i l i t y of t h e pneumatically powered s e c t i o n of t h e p r e s s u r e r e g u l a t o r valve t o remain closed during engine starts, leading t o long start t i m e s and poor engine a c c e l e r a t i o n c h a r a c t e r i s t i c s . This l a c k of pneumatic power w a s a r e s u l t of lower bleed duct p r e s s u r e s than preducted which are believed t o be caused by higher l o s s e s i n t h e engine bleed p o r t s and high-pressure ducting. T h i s problem w a s solved by p l a c i n g a motorized valve i n series w i t h t h e r e g u l a t o r v a l v e which ensures a p o s i t i v e c l o s u r e during low-speed engine operation.

Additional information on t h e engine operation and ground t e s t can be found i n r e f e r e n c e s 14-16.

1 4 F l i g h t Performance A s a p a r t of t h e management approach discussed earlier i n t h i s paper, r i g i d performance requirements were not imposed on Boeing, t h e a i r p l a n e prime c o n t r a c t o r . However, ambitious performance g o a l s w e r e e s t a b l i s h e d a t t h e start of t h e p r o j e c t and c u r r e n t p r e d i c t i o n s i n d i c a t e t h a t most of t h e s e w i l l b e m e t o r exceeded. The reason f o r t h e high-performance levels i s t o provide t h e QSRA w i t h t h e maximum amount of r e s e a r c h c a p a b i l i t y . P r o p u l s i v e - l i f t , r o l l a c c e l e r a t i o n , approach c a p a b i l i t y , and low community n o i s e are technology t a r g e t s t h a t w e r e emphasized by N A S A and t h a t were of primary importance i n t h e development of t h e QSRA design. Some of t h e more s i g n i f i c a n t performance c a p a b i l i t i e s of t h e QSRA are summarized i n t a b l e 4.

L i f t capability-One of t h e primary performance g o a l s , a minimum usable approach l i f t c o e f f i c i e n t (CL) of 4.6, is expected t o b e exceeded by about 16% a f t e r allowance has been made f o r commercial f l i g h t s a f e t y margins. Figure 21 compares t h e l i f t performance of t h e QSRA t o t h a t of a standard medium commer- c i a l j e t t r a n s p o r t ( t h e B-727) and t o t h e Boeing Advanced M i l i t a r y STOL Trans- p o r t (YC-14). The h i g h - l i f t c a p a b i l i t y f o r t h e QSRA w a s achieved by applying BLC, as discussed previously. This four- p r o p u l s i v e - l i f t and leading edge engine configuration permits a l a r g e r span USB f l a p and reduces adverse yaw and r o l l moments with one engine i n o p e r a t i v e , t h u s y i e l d i n g t h e improvement i n l i f t over t h e twin-engine YC-14. The a c t u a l f l i g h t d a t a i n d i c a t e t h a t t h e QSRA performance i s n e a r e r t o p r e d i c t i o n a t t h e higher angles of a t t a c k than t o t h e wind-tunnel data. The reasons f o r t h i s are complex, but are believed t o be due t o t h e f a c t t h a t t h e QSRA USB performance c o n f i g u r a t i o n has not been t r u l y optimized y e t and performance i s expected t o improve as more is known about t h e flow over t h e USB p o r t i o n of t h e wing.

QSRA research mission is Another performance area which is v i t a l t o t h e t h e r o l l c o n t r o l responsiveness of t h e a i r p l a n e . The QSRA r o l l a c c e l e r a t i o n This high (ZSRA i s compared t o t h a t of s e v e r a l o t h e r a i r p l a n e s i n f i g u r e 22.

r o l l c o n t r o l e f f e c t i v e n e s s is achieved by i n c o r p o r a t i n g blown a i l e r o n s , mini- mizing r o l l i n e r t i a , minimizing engine-out r o l l i n g moment, and by t h e assymet- r i c use of t h e double-slotted f l a p s f o r t r i m . This r o l l - c o n t r o l power is t h e research f l y i n g w i l l be done with one of t h e important because much of outboard ( c r i t i c a l ) engines shut down under unfavorable conditions i n order This curve has not been v e r i - t o develop criteria f o r f u t u r e STOL a i r c r a f t .

f i e d i n f l i g h t at t h i s t i m e but w i l l be explored during t h e next phase of f l i g h t t e s t i n g a t Ames.

STOL operating envelope-The STOL operating envelope of t h e QSRA with a l l engines operating is shown i n f i g u r e 23 and t h e envelope w i t h t h e c r i t i c a l engine ( e i t h e r outboard engine) i n o p e r a t i v e is shown i n f i g u r e 24. These l i m i - f i g u r e s show t h e aerodynamic c a p a b i l i t y of t h e a i r p l a n e . P i t c h c o n t r o l t a t i o n s of t h e present c o n f i g u r a t i o n prevent o p e r a t i o n a t f u l l USB f l a p d e f l e c - t i o n , at 100% t h r u s t , and a t low speeds. S i m i l a r l y , d i r e c t i o n a l c o n t r o l con- s i d e r a t i o n s l i m i t t h e minimum speed w i t h an engine out i n a go-around config- u r a t i o n . However, f u t u r e modifications t o t h e QSRA empennage, which are p r e s e n t l y under consideration, w i l l make it p o s s i b l e t o explore a l l corners of t h e envelope. The e x i s t i n g c o n f i g u r a t i o n can s a f e l y o p e r a t e at a lift c o e f f i c i e n t of 5.5 w h i l e maintaining speed, angle of a t t a c k , maneuver, and go-around climb margins.

When t h e t h r u s t i s increased t o 100% i n a go-around s i t u a t i o n , t h e USB f l a p s are autmoatically r e t r a c t e d t o t h e go-around s e t t i n g .

A s shown i n f i g u r e 24, t h i s permits a climb a n g l e of +2O (equivalent t o a rate of climb of 1.22 m / s (240 f t / m i n ) ) a t an approach l i f t c o e f f i c i e n t of 5.5 with t h e c r i t i c a l engine i n o p e r a t i v e . A s t h e speed is allowed t o i n c r e a s e , t h e climb angle i n c r e a s e s t o over +6O.

These o p e r a t i n g envelopes have been p a r t i a l l y v e r i f i e d i n t h e i n i t i a l Boeing f l i g h t test and a l s o i n t h e Ames f l i g h t test. Values of a i r p l a n e drag are s l i g h t l y higher than t h o s e p r e d i c t e d and performance at t h e very high f l a p s e t t i n g s (over 63') i n d i c a t e s t h a t t h e r e is some flow s e p a r a t i o n and s l i g h t l y lower t u r n i n g angles than w e r e a t t a i n e d i n wind-tunnel tests. However, as discussed earlier, it is believed t h a t configuration optimization and a b e t t e r understanding of p r o p u l s i v e - l i f t aerodynamics w i l l allow t h e a i r p l a n e t o even- t u a l l y exceed performance p r e d i c t i o n s a t t h e h i g h e s t STOL f l a p s e t t i n g s . This flow f i e l d h a s been explored t o a minor e x t e n t ( t u f t s ) during t h e f i r s t series of Ames tests and w i l l be explored i n g r e a t depth i n t h e next phase of t h e Ames f l i g h t r e s e a r c h program.

Approach angle-A s h o r t - f i e l d a i r p l a n e r e q u i r e s a s t e e p descent c a p a b i l i t y (high approach angle) i n order t o minimize t h e required a i r s p a c e i n t h e t e r m i - n a l area, as w e l l as t o minimize community n o i s e e f f e c t s . The USB nozzle and f l a p s of t h e QSRA have been designed t o provide e x c e p t i o n a l l y high flow turn- i n g of t h e engine exhaust, y i e l d i n g high l i f t approach l i f t c o e f f i c i e n t s (> 5.5) which enable t h i s a i r c r a f t t o achieve very s t e e p approaches with f u l l s a f e t y margins. Figure 25 g i v e s a comparison of t h e QSRA STOL c a p a b i l i t i e s and t h e descent angle and ground r o l l of a conventional t r a n s p o r t a i r c r a f t landing. A t t h e same d i s t a n c e from t h e a i r p o r t , t h e QSRA is more than t w i c e as high as t h e conventional a i r l i n e r and i t i s a b l e t o s t o p on t h e runway b e f o r e today's commercial t r a n s p o r t s complete t h e i r f l a r e and touch down on t h e runway. The landing and takeoff performance of t h e QSRA has been v e r i f i e d during t h e i n i t i a l f l i g h t test with ground r o l l d i s t a n c e of 202.4 m (664 f t ) during a maximum performance takeoff and of less than 167.6 m (550 f t ) during a STOL landing.

Because n o i s e a t t e n u a t e s r a p i d l y with d i s t a n c e , t h e higher approach a l t i - tude of t h e QSRA is a b i g f a c t o r i n reducing community n o i s e e f f e c t s ; t h i s can be increased even more by landing toward t h e c e n t e r of t h e runway.

height Another technique t h a t may reduce community n o i s e e f f e c t s , by keeping t h e is a s p i r a l o r n o i s e completely w i t h i n t h e a i r p o r t boundaries during t a k e o f f , c i r c l i n g approach and departure. Simulation s t u d i e s have shown t h a t t h e QSRA n o i s e can be confined t o t h e boundaries of a t y p i c a l g e n e r a l a v i a t i o n a i r p o r t , and f l i g h t test h a s shown t h a t t h e QSRA is capable of a 337.1-m (1106 f t ) r a d i u s d e p a r t u r e with a 30" bank angle with an i n c r e a s e i n a l t i t u d e of 884 m (2900 f t ) a f t e r a f u l l 360° t u r n .

Acoustic Performance One of t h e primary g o a l s of t h e QSRA program w a s t o have a 90-EPNdB community n o i s e impact area of no more than 2.5 km2 (1 mile2) f o r a 668,182-M (150,000 lb) commercial a i r p l a n e based on QSRA technology. Figure 26 shows how t h i s g o a l compares with t h e noise-impact area of a c u r r e n t medium short- h a u l commercial t r a n s p o r t (B-737, DC-9). The a c t u a l n o i s e levels of t h e QSRA were measured during t h e f i n a l phase of t h e Boeing f l i g h t - t e s t program, e x t r a - polated, and compared t o t h e program goals. The maximum e f f e c t i v e perceived n o i s e level (EPNL) measured on t h e 152.4-m (500 f t ) s i d e l i n e during takeoff w a s 93.5 EPNdB and t h e g o a l w a s 92 EPNdB; during landing it w a s 89 EPNdB and t h e goal w a s 90 EPNdB. As can be seen, t h e values are s l i g h t l y higher during takeoff and s l i g h t l y lower during landing. (It should be noted t h a t t h e s e takeoff and landing n o i s e levels are based on a 152.4-m (500 f t ) s i d e l i n e and hence t h e takeoff n o i s e is s u b s t a n t i a l l y g r e a t e r than t h a t which would be measured i n accordance with FAR 36.) S e v e r a l major d i f f e r e n c e s exist between .

t h e QSRA and any p o t e n t i a l t r a n s p o r t a i r c r a f t based on t h e QSRA technology, w i t h t h e most important of t h e s e being t h e high drag c o n f i g u r a t i o n of t h e QSRA during t a k e o f f , due t o t h e absence of f a i r i n g s and r e t r a c t a b l e landing gear, QSRA h a s n o n r e t r a c t a b l e v o r t e x g e n e r a t o r s and nozzle and t h e f a c t t h a t t h e doors i n t h e takeoff and c r u i s e configuration. The c l e a n t r a n s p o r t configura- t i o n would r e s u l t i n a takeoff/climbout speed i n c r e a s e from t h e 90 knot QSRA speed t o about 130 knots, r e s u l t i n g i n a decrease i n t h e takeoff s i d e l i n e n o i s e l e v e l t o approximately 91.5 EPNdB. Based on t h i s clean configuration QSRA, a 668,182-N (150,000 l b ) commercial t r a n s p o r t would have a 90 EPNdB n o i s e impact area of 7.03 km2 (2.8 mile2) compared t o t h e QSRA goal of 2.51 are preliminary r e s u l t s based on a km2 (1 m i l e 2 ) . These n o i s e e x t r a p o l a t i o n s l i m i t e d d a t a base and are believed t o be considerably l a r g e r than t h e n o i s e areas t h a t w i l l be achievable by a commercial t r a n s p o r t based on t h e QSRA technology. T h i s b e l i e f is based on a number of f a c t o r s , discussed i n t h e following s e c t i o n s , t h a t can be eliminated i n any f u t u r e t r a n s p o r t i f they are i s o l a t e d as s i g n i f i c a n t n o i s e sources (e.g., n o n r e t r a c t i n g v o r t e x g e n e r a t o r s ) .

Data analysis-The d a t a w e r e analyzed by t h e systems and methods used i n FAR-36 n o i s e c e r t i f i c a t i o n s , with 1 / 3 o c t i v e band s p e c t r a i n t e g r a t e d over 0.5- sec periods a t increments of 0.5 sec. Computer processing mated a c o u s t i c d a t a with t h e a i r p l a n e p o s i t i o n as determined o p t i c a l l y and with t h e f l i g h t p r o f i l e d a t a , s y n t h e s i z i n g f l y o v e r n o i s e t i m e h i s t o r i e s f o r t h e v a r i o u s r e f e r e n c e f l i g h t p r o f i l e s .

Far f i e l d results-The community n o i s e l e v e l d a t a r e s u l t from measurements made w i t h a p r e c i s i o n of about 21 EPNdB i n t h e EPNL measurements and about +2 PNdB i n PNL measurements. The n o i s e l e v e l s along t h e f l i g h t p a t h are higher with t h e USB flaps. r e t r a c t e d than they were with a 30° f l a p s e t t i n g and t h e A l - s i d e l i n e n o i s e l e v e l s w e r e r e l a t i v e l y unaffected by f l a p configuration.

though a i r f r a m e n o i s e w a s present i n some measurements, it d i d not s i g n i f i - The measured n o i s e exceeded c a n t l y i n f l u e n c e t h e PNL and EPNL n o i s e l e v e l s .

p r e d i c t i o n s , with t h e higher l e v e l s appearing t o b e r e l a t e d t o a random aero- dynamic n o i s e generated by i n t e r a c t i o n of turbulence w i t h i n t h e j e t flow w i t h It i s believed t h a t one p o s s i b l e source of t h i s n o i s e t h e wing t r a i l i n g edge.

I n a d d i t i o n , a l a r g e , low-frequency (200 Hz) may be the v o r t e x generators.

component seems t o be t h e r e s u l t of engine i n s t a l l a t i o n e f f e c t s .

An a d d i t i o n a l anomaly appeared i n some of t h e d a t a f o r r e t r a c t e d USB f l a p t h a t w a s taken a t a d i f f e r e n t t i m e than t h e rest of t h e d a t a . The levels of t h i s d a t a p o i n t appeared t o b e e s s e n t i a l l y t h e same as t h e 30° f l a p data.

This lower f a r - f i e l d n o i s e l e v e l w a s corroborated by near-field measurements discussed in the next section. There is no reasonable explanation of this anomaly at this time but subsequent testing of the QSRA will investigate this effect.

Near-field measurements-Near-field noise measurements were made with eight microphones flush-mounted on the fuselage exterior surface and four located inside the aircraft. These measurements showed that noise levels in- creased uniformly with engine power levels, approximating a 40 log V j rela- tionship, with maximum exterior noise levels of about 150 dB. The interior noise levels with inboard engines shut down were about 10 dB less than with all engines operating, especially in locations where flow attachment and fuse- lage scrubbing occurred. The maximum measured interior noise levels were 118 dB in the aft cabin at high power settings. It should be noted that the fuse- lage interior is untreated and is not representative of the noise levels that would exist in a similar commercial transport aircraft.

Data significance-The acoustic data presented in this section are pre- liminary and do not represent a complete analysis or a good data base. Rather, they represent a starting point from which to build a more complete understand- ing of propulsive-lift aircraft noise, and to develop the required flight ex- periments that will provide a technology base for future transports based on QSRA program goals. As has occurred in several other areas, the QSRA acoustic configuration has not been optimized. Optimization of the configuration can be expected to reduce the noise levels from those measured in these tests. It should be noted that a reduction in measured noise level of only 2 dB will result in noise impact area reduction of approximately 60%. Additional data on the QSRA flight tests and acoustic tests are given in reference 17.

FUTURE PLANS AND EXPERIMENTS Initial Tests and Configuration Optimization The initial NASA flight program at Ames Research Center will last for approximately 1% years and will be devoted to envelope documentation with con- figuration optimization as required. As discussed in several previous sec- tions, small changes in the configuration or flow field can have a significant effect on propulsive-lift aircraft performance. During these initial tests, a series of experiments will be performed to define the flow field and the effect of this flow field on the QSRA powered-lift performance. For example, a number of experiments will be performed to determine the effect of vortex generator size and location on the QSRA aerodynamic performance and far-field acoustics. Aerodynamic performance will be documented and compared to a ref- erence baseline as provided by a relatively sophisticated mathematical model of the QSRA and by the use of flow-visualization techniques such as tufts.

The acoustic measurements will be cross correlated with near-field measurements ' in order to isolate and identify the sources of the noise, and these experi- In addition, ments will be augmented with small scale tests as required.

alternative flight profiles will be evaluated in order to further reduce com- munity noise impact areas.

One interesting modification under consideration is the replacement of the present blown leading edge with an unblown leading edge slat. Wind-tunnel tests described earlier showed that an unblown leading edge would degrade per- formance by reducing the angle-of-attack margin 4' to 5 O (ref. 18). If this can be verified in flight, future designers of high-performance STOL aircraft will have a firm technical base for the selection of blown versus unblown lead- ing edges.

Another modification planned for the QSRA is to increase the gross weight to 267,273 N (60,000 l b ) . This will provide a wing loading of 4795 N/m2 (100 lb/ft2) and give the QSRA the capability of operating over a range of wing ' loadings from 3117 to 4795 N/m2 (65 to 100 lb/ft2) to increase its research versatility.

Flight Experiments After the initial NASA flight research program and configuration optimi- zation, the QSRA will be made available for the flight-experiments program.

In the initial sections of this paper, the concept of a research aircraft being a facility for flight research was discussed. When the QSRA enters the flight- experiments phase it will fulfill this goal and become a national facility for flight research. Research personnel within NASA are planning a program of flight experiments. Some of the experiments will be accomplished as in-house efforts; others will be done jointly with other government agencies, for exam- ple, the development of certification criteria for future STOL aircraft. In other cases, the work will be contracted, particularly when the experiment in- volves structural modification to the airplane or the development and instal- of new equipment. The QSRA is, however, a national flight facility.

lation As such, it is available to the aeronautical community in the same way that a NASA wind tunnel or simulator is available.

QSRA workshop-On November 29 and 30, 1978 a workshop will be held at Ames Research Center in order to provide industry, universities, and govern- ment agencies with information on the capabilities of the QSRA and to provide a mechanism by which participation in the flight experiments program can be implemented. It is hoped that this procedure will lead to broad participation by the aeronautical community in the QSRA flight research program.

Although the flight-experiments phase will not "officially" begin for several years, it is believed that many experiments, particularly, self- contained experiments, can be flown on the QSRA during the initial flight tests. In addition, some experiments, such as acoustic measurements and cor- relation of small-scale testing with the QSRA, can and should be done concur- rently with the early flight program. For example, one series of acoustic experiments which is under consideration involves the use of Ames' quiet noise measuring airplane, the YO-3A, to make free-field acoustic measurements of the QSRA flap and inlet noise. A number of other experiments are planned in vari- ous research areas such as avionics, computer control systems, inlet flow fields, acoustics, structural vibration, and aerodynamic performance.

Flight demonstration-Another activity in the early planning stages is demonstration flights at airports in the San Francisco Bay Area for potential users of the technology. This would include airline and airport officials and technical personnel from interested aircraft manufacturers.

The QSRA will not it is qualified only for a crew of two research carry passengers because pilots. However, flight demonstrations will expose potential users of this technology to the short-field capability, the maneuverability, and the low community noise levels of which the QSRA is capable. Late in the initial flight program, it is also planned that qualified pilots from other organiza- tions will be invited to fly and evaluate the QSRA with a NASA research pilot as an "instructor-pilot.'' This was done successfully with the Augmented Jet Flap STOL Research Airplane and will be repeated with the QSRA.

REFERENCES 1. Quigley, Hervey C.; Innis, Robert C.; and Grossmith, Seth: A Flight In- vestigation of the STOL Characteristics of an Augmented Jet Flap STOL Research Aircraft. NASA TM X-62334, 1974.

2. Quiet Propulsive Lift Research Aircraft Design Study. NASA CR-137557, 1 9 74.

3. Quiet Short Haul Research Aircraft Design Study. NASA CR-137554, 1974.

4. Cochrane, John A.; and Carros, Robert J.: Hybrid Upper Surface Blobm Flap Propulsive-Lift Concept for the Quiet Short-Haul Research Aircraft. AIAA Paper 75-1220, Oct. 1975.

5. Cochrane, John A.; and Boissevain, Alfred G.: Quiet Short-Haul Research Aircraft-Current Status and Future Plans. AIAA Paper 78-1468, Aug.

1978.

6 . Shovlin, Michael D.: Effects of Inlet Airframe Integration on the Inlet of a USB Four Engine STOL Airplane. AIAA Paper 78-959, July 1978.

7. Nickson, Theodore B . : Large Scale Wind Tunnel Investigation of the Quiet Short-Haul Research Aircraft (QSRA) Configuration. NASA CR-152095, 1978.

8 . Gunnarson, Daniel W.; and McArdle, Jack C . : Development and Test of an Inlet and Duct to Provide Airflow for a Wing Boundary Layer Control System. AIAA Paper 78-141, Jan. 1978.

9. Reshotko, Meyer; Karchmer, Allen N.; Penko, Paul F.; and McArdle, Jack G . : Core Noise Measurements on a YF-102 Turbofan Engine. AIAA Paper 77-21, Jan. 1977.

10. Middleton, Robie; and Vincent, James H . : Quiet Short-Haul Research Air- craft Phase I1 Flight Simulation Math Model-Final Report. NASA CR- 152197, 1978.

11. Wilcox, Darrell E.; and Quigley, Hervey C.: V/STOL A i r c r a f t Simulation- Requirements and C a p a b i l i t i e s at Ames Research Center. AIAA Paper 78-1515, Aug. 1978.

12. Riddle, Dennis W.: A P i l o t e d Simulator Analysis of t h e Carrier Landing of t h e Quiet Short-Haul Research A i r c r a f t . N A S A TM-58508, C a p a b i l i t y 1978.

13. Stevens, Fred: I Configuration D e f i n i t i o n Document of t h e QSRA. (Proposed N A S A high number c o n t r a c t o r r e p o r t . ) 14. M c N e i l l , John M.: QSRA/YF-102 Engine Performance and Control. Boeing Document D340-10206, Boeing Commercial Airplane Company, S e a t t l e , Wash- ington, 1978.

15. McNeill, John M.; and Harkonen, Dennis L.: QSRA Ground T e s t Report Volume I-Propulsion and Fuel System Testing. Boeing Document 0340-13701-1, Boeing Commercial Airplane Company, S e a t t l e , Washington, 1978.

16. Shah, Ani1 D.: QSRA Ground T e s t Report Volume II-Systems. Eoeing Docu- ment D340-13701-2, Boeing Commercial Airplane Company, S e a t t l e , Washing- ton, 1978.

17. Analysis of C o n t r a c t o r ' s Taxi and F l i g h t T e s t of t h e QSRA. (Proposed N A S A high number c o n t r a c t o r r e p o r t . ) 18. Holtman, Donald N.; and Howard, W e s M.: Large S c a l e Wind Tunnel I n v e s t i - g a t i o n f o r Future Modifications t o t h e Quiet Short-Haul Research A i r - c r a f t (QSRA) Configuration. (Proposed N A S A high number c o n t r a c t o r r e p o r t . ) TABLE 1.-QSRA INITIAL GOALS AND REQUIREMENTS (PARTIAL LIST) Requirements Day, VFR operation only Crew of two research pilots 0 Modification of GFE C-8A Buffalo 0 Use of four GFE YF-102 engines 0 Hybrid upper surface blowing propulsive lift concept Goals

0 Approach lift coefficient - 4.6 (steep approach with margins)

0 Approach path of -7.5' with margin for gusts, wind, etc.

90 EPNdB combined takeoff and landing footprint area, when scaled to 668,182 N (150,000 lb) of 2.5 km2 (1 mile2) Minimum duration of test mission-50 min e Minimum wing loading at gross weight = 3117 N/m2 (65 lb/ft2) 0 Maximum cruise speed 160 knots 0 Wing/nacelle configuration representative of cruise at M = 0.74.

TABLE 2. -QSRA FUNDING DISTRIBUTION Preliminary design s t u d i e s $ 2,000,000 Wind t u n n e l and engine tests 1,000,000 Engine program 2,000,000 Airplane d e t a i l design and f a b r i c a t i o n 22,000,000 Proof-of-concept f l i g h t test 2,000,000 T o t a l funding a v a i l a b l e $29,000,000 TABLE 3.-BLC F L O W DISTRIBUTION Engine p o s i t i o n BLC segment 1 Right a i l e r o n 2 Right leading edge 3 L e f t leading edge 4 L e f t a i l e r o n TABLE 4.- QSRA PREDICTED PERFORMANCE Approach l i f t c o e f f i c i e n t ( s t e e p approach with margins) 5.5 Approach path with margin f o r wind, g u s t s , etc. -7.5" Duration of STOL test mission 102 min 426.7 m (1400 f t ) Landing f i e l d - l e n g t h a t 213,370 N (48,000 l b ) (W/S = 3836 N/m2 (80 l b / f t 2 ) ) (1.67 f a c t o r over 10.7-m (35 f t ) o b s t a c l e ) Takeoff f i e l d l e n g t h a t 668,182 N (150,000 l b ) 403.9 m (1325 f t ) (10.7-m (35 f t ) o b s t a c l e with c r i t i c a l engine i n o p e r a t i v e (CEI) a t d e c i s i o n speed) Turn r a d i u s a t 30" bank a n g l e 213.4 m (700 f t )

Figure 1 . - The Quiet Short-Haul Research Airplane (QSRA) Performing a S ' N X

approach p r i o r to landing a t Ames Research Center.

Figure 2.-The 0.55-scale QSRA model mounted in the Ames 40- by 80-Foot Wind Tunnel.

Figure 3.-The YF-102 (QSRA) engine i n s t a l l e d i n t h e L e w i s Vertical L i f t Fan F a c i l i t y i n preparation f o r confluent flow b a s e l i n e t e s t i n g .

Figure 4 . - P i l o t ' s view from t h e cab of t h e Ames FSAA during t h e QSRA f l i g h t simulation.

TOGW. Ib 5 o . m 222260N WING AREA, ftz 6 ~ ) 65.74m2 WING LOADING, lbIh2 83 3987.4Nlm2 ASPECT RATIO 9 THRUSTlWEIGHT . H I

/+ 73.5 h A 93.25 R

22.40 m 28.42 m Figure 5. -The QSRA a i r p l a n e layout.

Figure 6.-The QSRA wing and fuselage under construction a t t h e Boeing Developmental Center i n Seattle, Washington.

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 Figure 7 . - A cutaway view of t h e YF-102 (QSRA) engine which w a s b u i l t by AVCO-Lycoming Division.

T

18.56 i n .

0.471 m

I

N, 7600 (FAN) (POWER TURBINE 17600) N, 19660 SFC 0 . 4 1 MGT 1665°F 907'C MAX. THRUST 7500 Ib 33409.1 N WEIGHT DRY 1215 Ib 5 5 1 . 1 kg TOTAL AIRFLOW 267 Ib/sec 1 2 1 . 1 kg/sec CORE AIRFLOW 37 Iblsec 1 6 . 8 kg/sec B.P.R. 6 . 2 Figure 8. -YF-102 (QSRA) engine layout.

I I 6' INBD 4.6" OUTBD 134.30 in.

k p 3.411 r n - 4

Figure 9 . - T h e QSRA nacelle layout.

STRUCTURAL COWL AND NOZZLE ASSY ACCESS DOOR TO FWD ACCESSORIES STRUCTURAL FAN DUCT RAMP AND UPPER SUPPORT LOWER FAN DUCT REMOVABLE

\ INBOARD, OUTBOARD

LOWER COWL DOOR

/F=-

COWL WING FAIRING INBOARD, OUTBOARD Figure l O . - V i e w of the QSRA nacelle main structural elements.

ORIGINAL PAGE ‘D

OF POOR QUALITY NOSE COWL PRIMARY NOZZLE COOLING MAN1FOLD ACCESSORY DRIVE SYSTEM Figure 11.-The engine build up showing the YF-102 (QSRA) engine i n s t a l l e d i n an assembly consisting of the engine, core cowl, i n l e t , and primary nozzle.

EXHAUST MIXING

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7 PNdB INLET NOISE ATTENUATION DOUBLE LAYER LINING WING SHIELDING 12 PNdB AFT FAN ATTENUATION SINGLE LAYER LINING INNER AND OUTER FAN DUCT WALLS Figure 12.-Location of the acoustic lining panels i n the QSRA i n l e t and a f t fan f l o w passages.

HIGH-PRESSURE Figure 13.-The layout of t h e QSRA boundary-layer c o n t r o l (BLC) system.

The two phantom l i n e ducts only connect t h e leading and t r a i l i n g edge systems i n t h e event of an engine f a i l u r e .

BLC EJECTOR FIRE BOTTLES DRAIN MAST STARTER HYD PUMP CDS GEN CDS COOLER AND FAN CDS ACCUMULATOR CDS FILTER Figure 14.-Systems layout i n t h e QSRA nacelle.

ORIGINAL PAGB 18

OF POOR QUA1,ITY

NOZZLE n 2 o O O r z AIR SUPPLY TO

-

E A BLC SYSTEM m t

t t HIGH - PRESSURE

U 4 0 0 ~ 300 2 AIR IN m

-

3 1200 I-- v) I c a (1 I

f 800-

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m z

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400- m t; MAX. ENGINE THRUST z 4 * 0 - Figure 1 5 . - T h e performance of t h e QSRA a i l e r o n BLC system shown as a f u n c t i o n The BLC e j e c t o r which combines t h e f a n and core a i r from t h e engine t h r u s t .

engine t o provide t h e BLC system a i r flow is shown i n t h e i n s e r t .

ELEVATOR (DOUBLE SLOTTED) SPO I LE RS RUDDER (DOUBLE HINGE) USB FLAPS &FLAP (DOUBLE SLOTTED) A AILERON Figure 16. -Location of t h e QSRA main f l i g h t c o n t r o l s .

100 120 E e

g 100

c C

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f!

s

W 8 70 60 I - - v) W K a 0 60 40

E

P

ESTRICTED ZONE I-

3 50 20

K K

s

40 0 20 30 40 50 60 70 80 90 100 CORRECTED FAN SPEED, N p / a , % rpm Figure 17.-The r e l a t i o n s h i p of t h e f a n and c o r e compressor speeds w i t h each o t h e r and w i t h engine t h r u s t level f o r t h e YF-102 (QSRA) engine.

SEA LEVEL STATIC 3% FAN BLEED 30 HPX OUTBOARD ENGINE

-----

INBOARD ENGINE

-

32,000 HIGH SHAFT 7 0 0 0 r / T R Q U E LIMIT

6500 E====- +. \

-

P I - - 28,000 I-- v) ED v) K 6000- I I I- c

P

-I -I

2 5500-

-

t; 24,000

G

z

-

20.000 4500 \

0 20 40 60 80 100 120 AMBIENT TEMPERATURE, TAM, O F I I I I I I I I -20 -10 0 10 20 30 40 50 "C Figure 18.- The effect of ambient temperature on t h e YF-102 (QSRA) engine t h r u s t . The engine is torque-limited at t h e lower temperatures and environ- ment c o n t r o l system (ECS) bleed is only taken from t h e inboard engines.

YF-102 FAN 0 ENG. 1 (YFO1) BYPASS MAP A ENG-2 (YF02) ORIGINAL PAGE I$ N E3 ENG-3 (YF03)

OF POOR QUALITY

loo%-$= 7600rpm 0 ENG.4(YF05) RELATIVE NOZZLE AREA FROM UNTRIMMED NOZZLE 1.45 SURGE LINE, bp 1-40 UPPER LIMIT FOR ENGINE

2 1.35

a 1 . 3 0 ~ ACCEPTANCE A-9 PROGRAM* TESTS DURING 1.25 a 3 1.20 ROTATING STALL LINE,

E 1.15

" n

z 1.10 ESTIMATED STEADY STATE OPERATING LINES (S.L.S.)

' 1.051 I I

40 60 80 100 120 140 160 180 200 220 CORRECTED BYPASS AIRFLOW,m, Ibfsec 1 I I I I I I I I 20 30 40 50 60 70 80 90 100 kd= Figure 19.-The YF-102 (QSRA) engine fan map. Results of the i n s t a l l e d engine ground t e s t s , which indicate that there is adequate operating margin for each engine, are shown.

e- ENG .3 /-- ENG. 2 ENG *4 BLC BLEED "ON" STD DAY S.L.S.

I I I I I I 0 2 4 6 8 10 12 TIME, sec Figure 20.-The acceleration t i m e history for each of the QSRA engines.

Note the e f f e c t of the s l i g h t l y out-of-phase BLC valve operation on the acceleration of engine No. 2.

0 TRIMMED LIFT BASED ON WIND-TUNNEL DATA 0 FLAPS IN APPROACH ATTITUDE 0 ALL ENGINES OPERATING

7l 6

0 , 5 10 15 20 25 30

ANGLE OF ATTACK, deg

-

1.4

-

1.2 \ U E -

. 1.0

-

d

w A w

-

0 .6 A .J OTHER STOL AIRPLANES

g . 4 -

A --W 0 1 I I I I I I I 2 2.5 3 3 . 5 4 4.5 5 5.5 APPROACH LIFT COEFFICIENT, C L ~ ~ ~ Figure 22.- The predicted QSRA roll performance compared with previous STOL airplanes.

-3 6 ORIGINAL PAGE 'SS

OF POOR QUALlV

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-2 cs) 3 - 4 - G -6 - - -8

-

-10 Ag = 0.25

A a = l 5 0 - . I . \-

-12 1 I I I I I I --* I

1 2 3 4 5 6 7 8 9 CL Figure 23. -Approach performance of the QSRA showing flight safety margins.

B O s -2 -4 -6 -8 1 2 3 4 5 6 7 CL Figure 24. -Performance of the QSRA after a critical (inboard) engine has failed and the QSRA is reconfigured for go-around.

100 n APPROACH LANDING 1 ° ~ , ALTITUDE AT - I rnile(1- r n ~

,L\

\

RUNWAY THRESHOLD

0 I I I I , \

-6000 -5000 -4000 h I I I I I I I -2000 -1500 -1000 -500 0 500 1000 APPROACH AND LANDING DISTANCE, m Figure 25.-The STOL landing c a p a b i l i t y of t h e QSRA compared t o a conventional approach and landing.

QSRA (SCALED TO MEDIUM MEDIUM TRANSPORT TRANSPORT SIZE) , 1 I I I I 1 90 EPNdB FOOTPRINTS Figure 26.-A comparison of t h e n o i s e impact areas of a c u r r e n t medium trans- p o r t a i r c r a f t and of t h e same s i z e t r a n s p o r t which m e e t s t h e QSRA n o i s e goals. There is approximately an 80 t o 90 percent reduction i n community noise impact.

Ames Research Center, NASA Moffett Field, Calif. 94035 National Aeronautics and Space Administration 15. Supplementary Notes 16. Abstract An overview of the Quiet Short-Haul Research Aircraft (QSRA) Program is presented, with special emphasis on its propulsion and acoustic aspects.

The historical background and management approach which led to a success- ful, low-cost research aircraft, capable of high levels of low-speed per- formance and very low community noise impact, are discussed. A description of the NASA technical participation in the program including wind-tunnel testing, engine ground tests, and advanced aircraft simulation is given.

The aircraft and its systems are described and, measured performance, Although most program goals where available, is compared to program goals.

were met or exceeded, preliminary data indicate that additional research and development are needed in some areas of which acoustics is an example, Some of these additional research areas and potential experiments using the QSRA to develop this technology are discussed. The concept of the QSRA as a national flight-research facility is explained and future plans and programs utilizing it for this 'purpose are described.

7. Key Words (Suggested by AuthorlsJt 18. Distribution Statfment Short-haul transports Unlimited STOL aircraft Aircraft design

STAR Category - 05

this mJ 21. No. of P~QS-S 22. Rice.

20. Security Classif. 1 9. k u r i t v Classif. (of this report)

Unclassified Unclassified I 41 I $4.00

*For sale by the National Technical Information Service, Springfield, Virginia 22161

Overview

Key points

Document

An Overview of the Quiet Short-Haul

Research Aircraft Program

Michael D. Shovlin and John A. Cochrane

22 .4 m (73.5 ft), a wing area of 55.74 m2 (600 ft >, and a quarter chord sweep

0 Approach lift coefficient - 4.6 (steep approach with margins)

Figure 1 . - The Quiet Short-Haul Research Airplane (QSRA) Performing a S ' N X

/+ 73.5 h A 93.25 R

T

I

k p 3.411 r n - 4

\ INBOARD, OUTBOARD

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STAR Category - 05

Unclassified Unclassified I 41 I $4.00

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